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Entry, Descent, and Landing: 2000-2004 

This custom bibliography from the NASA Scientific and Technical Information Program lists a sampling of records 
found in the NASA Aeronautics and Space Database. The scope of this topic includes technologies for precision 
targeting and landing on "high-g" and "low-g" planetary bodies. This area of focus is one of the enabling 
technologies as defined by NASA's Report of the President's Commission on Implementation of United States 
Space Exploration Policy, published in June 2004. 



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Entry, Descent, and Landing 
2000-2004 



A Custom Bibliography From the 

NASA Scientific and Technical Information Program 



October 2004 



Entry, Descent, and Landing: 2000-2004 

This custom bibliography from the NASA Scientific and Technical Information Program lists a sampling of records 
found in the NASA Aeronautics and Space Database. The scope of this topic includes technologies for precision 
targeting and landing on "high-g" and "low-g" planetary bodies. This area of focus is one of the enabling 
technologies as defined by NASA's Report of the President's Commission on Implementation of United States 
Space Exploration Policy, published in June 2004. 

OCTOBER 2004 

20040120953 Computer Sciences Corp., Huntsville, AL, USA 
Atmospheric Models for Aerocapture Systems Studies 

Justus, C. G.; Duvall, Aleta; Keller, Vernon W.; December 19, 2003; In English, 16-19 Aug. 2004, Providence, RI, USA 
Contract(s)/Grant(s): NAS8-60000; No Copyright; Avail: Other Sources 

Aerocapture uses atmospheric drag to decelerate into captured orbit from interplanetary transfer orbit. This includes 
capture into Earth orbit from, for example, Lunar-return or Mars-return orbit. Eight Solar System destinations have sufficient 
atmosphere for aerocapture to be applicable - three of the rocky planets (Venus, Earth, and Mars), four gas giants (Jupiter, 
Saturn, Uranus, and Neptune), and Saturn's moon Titan. These destinations fall into two groups: (1) The rocky planets, which 
have warm surface temperatures (approx. 200 to 750 K) and rapid decrease of density with altitude, and (2) the gas giants and 
Titan, which have cold temperatures (approx. 70 to 170 K) at the surface or 1-bar pressure level, and slow rate of decrease 
of density with altitude. The height variation of average density with altitude above 1-bar pressure level for the gas giant 
planets is shown. The periapsis density required for aerocapture of spacecraft having typical values of ballistic coefficient (a 
measure of mass per unit cross-sectional area) is also shown. The aerocapture altitudes at the gas giants would typically range 
from approx. 150 to 300 km. Density profiles are compared for the rocky planets with those for Titan and Neptune. 
Aerocapture at the rocky planets would occur at heights of approx. 50 to 100 km. For comparison, typical density and altitudes 
for aerobraking operations (circularizing a highly elliptical capture orbit, using multiple atmospheric passes) are also indicated. 
Author (revised) 
Atmospheric Models; Aerocapture; Aerobraking; Planetary Atmospheres 

20040120944 NASA Langley Research Center, Hampton, VA, USA 

Preliminary Convective-Radiative Heating Environments for a Neptune Aerocapture Mission 

Hollis, Brian R.; Wright, Michael J.; Olejniczak, Joseph; Takashima, Naruhisa; Sutton, Kenneth; Prabhu, Dinesh; [2004]; In 
English; AI A A Atmospheric Flight Mechanics Conference and Exhibit, 16-19 Aug. 2004, Providence, RI, USA 
Contract(s)/Grant(s): NAS2-99092; NAS1-00135; NCC1-02043; 320-10-00 
Report No.(s): AIAA Paper 2004-5177; No Copyright; Avail: CASI; A03, Hardcopy 

Convective and radiative heating environments have been computed for a three-dimensional ellipsled configuration which 
would perform an aerocapture maneuver at Neptune. This work was performed as part of a one-year Neptune aerocapture 
spacecraft systems study that also included analyses of trajectories, atmospheric modeling, aerodynamics, structural design, 
and other disciplines. Complementary heating analyses were conducted by separate teams using independent sets of 
aerothermodynamic modeling tools (i.e. Navier-Stokes and radiation transport codes). Environments were generated for a 
large 5.50 m length ellipsled and a small 2.88 m length ellipsled. Radiative heating was found to contribute up to 80% of the 
total heating rate at the ellipsled nose depending on the trajectory point. Good agreement between convective heating 
predictions from the two Navier-Stokes solvers was obtained. However, the radiation analysis revealed several uncertainties 
in the computational models employed in both sets of codes, as well as large differences between the predicted radiative 
heating rates. 
Author 
Aerocapture; Convective Heat Transfer; Radiative Heat Transfer; Neptune Atmosphere; Aerothermodynamics 

20040120869 Morgan Research Corp., Huntsville, AL, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA 
Atmospheric Models for Aeroentry and Aeroassist 

Justus, C. G; Duvall, Aleta; Keller, Vernon W.; June 15, 2004; In English, 23-26 Aug. 2004, Moffett Field, CA, USA; Original 

contains color illustrations 

Contract(s)/Grant(s): NNM04AA02C; No Copyright; Avail: CASI; A02, Hardcopy 

1 



Eight destinations in the Solar System have sufficient atmosphere for aeroentry, aeroassist, or aerobraking/aerocapture: 
Venus, Earth, Mars, Jupiter, Saturn, Uranus, and Neptune, plus Saturn's moon Titan. Engineering-level atmospheric models 
for Earth, Mars, Titan, and Neptune have been developed for use in NASA s systems analysis studies of aerocapture 
applications. Development has begun on a similar atmospheric model for Venus. An important capability of these models is 
simulation of quasi-random perturbations for Monte Carlo analyses in developing guidance, navigation and control algorithms, 
and for thermal systems design. Characteristics of these atmospheric models are compared, and example applications for 
aerocapture are presented. Recent Titan atmospheric model updates are discussed, in anticipation of applications for trajectory 
and atmospheric reconstruct of Huygens Probe entry at Titan. Recent and planned updates to the Mars atmospheric model, in 
support of future Mars aerocapture systems analysis studies, are also presented. 
Author 
Atmospheric Models; Solar System; Aeroassist; Aerobraking; Aerocapture; Guidance (Motion) 



20040111219 NASA Langley Research Center, Hampton, VA, USA 
Structural Design for a Neptune Aerocapture Mission 

Dyke, R. Eric; Hrinda, Glenn A.; [2004]; In English, 16-19 Aug. 2004, Providence, RI, USA 

Contract(s)/Grant(s): 23-800-90-40 

Report No.(s): AIAA Paper 2004-5179; No Copyright; Avail: CASI; A03, Hardcopy 

A multi-center study was conducted in 2003 to assess the feasibility of and technology requirements for using aerocapture 
to insert a scientific platform into orbit around Neptune. The aerocapture technique offers a potential method of greatly 
reducing orbiter mass and thus total spacecraft launch mass by minimizing the required propulsion system mass. This study 
involved the collaborative efforts of personnel from Langley Research Center (LaRC), Johnson Space Flight Center (JSFC), 
Marshall Space Flight Center (MSFC), Ames Research Center (ARC), and the Jet Propulsion Laboratory (JPL). One aspect 
of this effort was the structural design of the full spacecraft configuration, including the ellipsled aerocapture orbiter and the 
in-space solar electric propulsion (SEP) module/cruise stage. This paper will discuss the functional and structural requirements 
for each of these components, some of the design trades leading to the final configuration, the loading environments, and the 
analysis methods used to ensure structural integrity. It will also highlight the design and structural challenges faced while 
trying to integrate all the mission requirements. Component sizes, materials, construction methods and analytical results, 
including masses and natural frequencies, will be presented, showing the feasibility of the resulting design for use in a Neptune 
aerocapture mission. Lastly, results of a post-study structural mass optimization effort on the ellipsled will be discussed, 
showing potential mass savings and their influence on structural strength and stiffness 
Author 
Aerocapture; Neptune (Planet); Structural Design; Space Missions; Spacecraft Design 



20040111218 NASA Langley Research Center, Hampton, VA, USA 
Aerocapture Performance Analysis for a Neptune-Triton Exploration Mission 

Starr, Brett R.; Westhelle, Carlos H.; Masciarelli, James P.; [2004]; In English, 16-19 Aug. 2004, Providence, RI, USA 

Contract(s)/Grant(s): 23-800-90-50 

Report No.(s): AIAA Paper 2004-4955; No Copyright; Avail: CASI; A03, Hardcopy 

A systems analysis has been conducted for a Neptune-Triton Exploration Mission in which aerocapture is used to capture 
a spacecraft at Neptune. Aerocapture uses aerodynamic drag instead of propulsion to decelerate from the interplanetary 
approach trajectory to a captured orbit during a single pass through the atmosphere. After capture, propulsion is used to move 
the spacecraft from the initial captured orbit to the desired science orbit. A preliminary assessment identified that a spacecraft 
with a lift to drag ratio of 0.8 was required for aerocapture. Performance analyses of the 0.8 L/D vehicle were performed using 
a high fidelity flight simulation within a Monte Carlo executive to determine mission success statistics. The simulation was 
the Program to Optimize Simulated Trajectories (POST) modified to include Neptune specific atmospheric and planet models, 
spacecraft aerodynamic characteristics, and interplanetary trajectory models. To these were added autonomous guidance and 
pseudo flight controller models. The Monte Carlo analyses incorporated approach trajectory delivery errors, aerodynamic 
characteristics uncertainties, and atmospheric density variations. Monte Carlo analyses were performed for a reference set of 
uncertainties and sets of uncertainties modified to produce increased and reduced atmospheric variability. For the reference 
uncertainties, the 0.8 L/D flatbottom ellipsled vehicle achieves 100% successful capture and has a 99.87 probability of 
attaining the science orbit with a 360 m/s V budget for apoapsis and periapsis adjustment. Monte Carlo analyses were also 
performed for a guidance system that modulates both bank angle and angle of attack with the reference set of uncertainties. 



An alpha and bank modulation guidance system reduces the 99.87 percentile DELTA V 173 m/s (48%) to 187 m/s for the 

reference set of uncertainties. 

Author 

Aerocapture; Neptune (Planet); Space Missions; Triton; Systems Analysis; Spacecraft Performance; Space Exploration 

20040111217 NASA Langley Research Center, Hampton, VA, USA 
Neptune Aerocapture Systems Analysis 

Lockwood, Mary Kae; [2004]; In English, 16-19 Aug. 2004, Providence, RI, USA 

Contract(s)/Grant(s): 23-800-90-10 

Report No.(s): AIAA Paper 2004-4951; No Copyright; Avail: CASI; A03, Hardcopy 

A Neptune Aerocapture Systems Analysis is completed to determine the feasibility, benefit and risk of an aeroshell 
aerocapture system for Neptune and to identify technology gaps and technology performance goals. The high fidelity systems 
analysis is completed by a five center NASA team and includes the following disciplines and analyses: science; mission 
design; aeroshell configuration screening and definition; interplanetary navigation analyses; atmosphere modeling; 
computational fluid dynamics for aerodynamic performance and database definition; initial stability analyses; guidance 
development; atmospheric flight simulation; computational fluid dynamics and radiation analyses for aeroheating environment 
definition; thermal protection system design, concepts and sizing; mass properties; structures; spacecraft design and 
packaging; and mass sensitivities. Results show that aerocapture can deliver 1.4 times more mass to Neptune orbit than an 
all-propulsive system for the same launch vehicle. In addition aerocapture results in a 3-4 year reduction in trip time compared 
to all-propulsive systems. Aerocapture is feasible and performance is adequate for the Neptune aerocapture mission. Monte 
Carlo simulation results show 100% successful capture for all cases including conservative assumptions on atmosphere and 
navigation. Enabling technologies for this mission include TPS manufacturing; and aerothermodynamic methods and 
validation for determining coupled 3-D convection, radiation and ablation aeroheating rates and loads, and the effects on 
surface recession. 
Author 
Aerocapture; Neptune (Planet); Systems Analysis; Technology Utilization; Aeroshells 

20040095913 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA, USA 
Mars Exploration Rovers Landing Dispersion Analysis 

Knocke, Philip C; Wawrzyniak, Geoffrey G.; Kennedy, Brian M.; Desai, Prasun N.; Parker, Timothy J.; Golombek, Matthew 

P.; Duxbury, Thomas C; Kass, David M.; [2004]; In English 

Report No.(s): AIAA Paper 2004-5093; No Copyright; Avail: CASI; A03, Hardcopy 

Landing dispersion estimates for the Mars Exploration Rover missions were key elements in the site targeting process and 
in the evaluation of landing risk. This paper addresses the process and results of the landing dispersion analyses performed 
for both Spirit and Opportunity. The several contributors to landing dispersions (navigation and atmospheric uncertainties, 
spacecraft modeling, winds, and margins) are discussed, as are the analysis tools used. JPL's MarsLS program, a 
MATLAB-based landing dispersion visualization and statistical analysis tool, was used to calculate the probability of landing 
within hazardous areas. By convolving this with the probability of landing within flight system limits (in-spec landing) for 
each hazard area, a single overall measure of landing risk was calculated for each landing ellipse. In-spec probability contours 
were also generated, allowing a more synoptic view of site risks, illustrating the sensitivity to changes in landing location, and 
quantifying the possible consequences of anomalies such as incomplete maneuvers. Data and products required to support 
these analyses are described, including the landing footprints calculated by NASA Langley' s POST program and JPL's AEPL 
program, cartographically registered base maps and hazard maps, and flight system estimates of in-spec landing probabilities 
for each hazard terrain type. Various factors encountered during operations, including evolving navigation estimates and 
changing atmospheric models, are discussed and final landing points are compared with approach estimates. 
Author 
Mars Exploration; Landing Modules; Roving Vehicles; Statistical Analysis; Land 

20040095912 NASA Langley Research Center, Hampton, VA, USA, Jet Propulsion Lab., California Inst, of Tech., Pasadena, 

CA, USA 

Mars Exploration Rovers Entry, Descent, and Landing Trajectory Analysis 

Desai, Prasun N.; Knocke, Philip C; August 11, 2004; In English, 16-19 Aug. 2004, Providence, RI, USA 

Contract(s)/Grant(s): 23-749-30-00 

Report No.(s): AIAA Paper 2004-5092; No Copyright; Avail: CASI; A02, Hardcopy 



The Mars Exploration Rover mission successfully landed two rovers 'Spirit' and 'Opportunity' on Mars on January 4th 
and 25th of 2004, respectively. The trajectory analysis performed to define the entry, descent, and landing (EDL) scenario is 
described. The entry requirements and constraints are presented, as well as uncertainties used in a Monte Carlo dispersion 
analysis to statistically assess the robustness of the entry design to off-nominal conditions. In the analysis, six-degree-of- 
freedom and three-degree-of-freedom trajectory results are compared to assess the entry characteristics of the capsule. 
Comparison of the preentry results to preliminary post-landing reconstruction data shows that all EDL parameters were within 
the requirements. In addition, the final landing position for both 'Spirit' and 'Opportunity' were within 15 km of the predicted 
landing location. 
Author 
Trajectory Analysis; Mars Landing; Roving Vehicles; Position (Location) 

20040086474 NASA Langley Research Center, Hampton, VA, USA 
Aeroassist Technology Planning for Exploration 

Munk, Michelle M.; Powell, Richard W.; [2000]; In English 

Report No.(s): AAS-00-169; No Copyright; Avail: CASI; A03, Hardcopy 

Now that the International Space Station is undergoing assembly, NASA is strategizing about the next logical exploration 
strategy for robotic missions and the next destination for humans. NASA's current efforts are in developing technologies that 
will both aid the robotic exploration strategy and make human flight to other celestial bodies both safe and affordable. One 
of these enabling technologies for future robotic and human exploration missions is aeroassist. This paper will (1) define 
aeroassist, (2) explain the benefits and uses of aeroassist, and (3) describe a method, currently used by the NASA Aeroassist 
Working Group, by which widely geographically distributed teams can assemble, present, use, and archive technology 
information. 
Author 
Aeroassist; International Space Station; NASA Space Programs; Space Exploration; Technological Forecasting 

20040085958 Morgan Research Corp., Huntsville, AL, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA 
Atmospheric Models for Aerocapture 

Justus, C. G.; Duvall, Aleta L.; Keller, Vernon W.; April 09, 2004; In English, 11-14 Jul. 2004, Fort Lauderdale, FL, USA 
Contract(s)/Grant(s): NNM04AA02C; No Copyright; Avail: CASI; A02, Hardcopy 

There are eight destinations in the solar System with sufficient atmosphere for aerocapture to be a viable aeroassist option 
- Venus, Earth, Mars, Jupiter, Saturn and its moon Titan, Uranus, and Neptune. Engineering-level atmospheric models for four 
of these targets (Earth, Mars, Titan, and Neptune) have been developed for NASA to support systems analysis studies of 
potential future aerocapture missions. Development of a similar atmospheric model for Venus has recently commenced. An 
important capability of all of these models is their ability to simulate quasi-random density perturbations for Monte Carlo 
analyses in developing guidance, navigation and control algorithm, and for thermal systems design. Similarities and 
differences among these atmospheric models are presented, with emphasis on the recently developed Neptune model and on 
planned characteristics of the Venus model. Example applications for aerocapture are also presented and illustrated. Recent 
updates to the Titan atmospheric model are discussed, in anticipation of applications for trajectory and atmospheric reconstruct 
of Huygens Probe entry at Titan. 
Author 
Atmospheric Models; Aerocapture; Huygens Probe; Environmental Monitoring 

20040085708 NASA Langley Research Center, Hampton, VA, USA 
Aeroheating Thermal Analysis Methods for Aerobraking Mars Missions 

Amundsen, Ruth M.; Dec, John A.; George, Benjamin E.; [2002]; In English; No Copyright; Avail: CASI; A03, Hardcopy 
Mars missions often employ aerobraking upon arrival at Mars as a low-mass method to gradually reduce the orbit period 
from a high-altitude, highly elliptical insertion orbit to the final science orbit. Two recent missions that made use of 
aerobraking were Mars Global Surveyor (MGS) and Mars Odyssey. Both spacecraft had solar arrays as the main aerobraking 
surface area. Aerobraking produces a high heat load on the solar arrays, which have a large surface area exposed to the airflow 
and relatively low mass. To accurately model the complex behavior during aerobraking, the thermal analysis must be tightly 
coupled to the flight mechanics, aerodynamics, and atmospheric modeling efforts being performed during operations. To 
properly represent the temperatures prior to and during the drag pass, the model must include the orbital solar and planetary 
heat fluxes. The correlation of the thermal model to flight data allows a validation of the modeling process, as well as 



information on what processes dominate the thermal behavior. This paper describes the thermal modeling method that was 

developed for this purpose, as well as correlation for two flight missions, and a discussion of improvements to the 

methodology. 

Author 

Aerobraking; Mars Missions; Elliptical Orbits; Thermal Analysis; Aerodynamic Heating 

20040068067 Computer Sciences Corp., Huntsville, AL, USA 

Connecting Atmospheric Science and Atmospheric Models for Aerocaptured Missions to Titan and the Outer Planets 

Justus, C. G.; Duvall, Aleta; Keller, Vernon W.; December 19, 2003; In English, 25-30 Apr. 2004, Nice, France 
Contract(s)/Grant(s): NAS8-60000; No Copyright; Avail: Other Sources; Abstract Only 

Many atmospheric measurement systems, such as the sounding instruments on Voyager, gather atmospheric information 
in the form of temperature versus pressure level. In these terms, there is considerable consistency among the mean atmospheric 
profiles of the outer planets Jupiter through Neptune, including Titan. On a given planet or on Titan, the range of variability 
of temperature versus pressure level due to seasonal, latitudinal, and diurnal variations is also not large. However, many 
engineering needs for atmospheric models relate not to temperature versus pressure level but atmospheric density versus 
geometric altitude. This need is especially true for design and analysis of aerocapture systems. Aerocapture drag force 
available for aerocapture is directly proportional to atmospheric density. Available aerocapture 'corridor width' (allowable 
range of atmospheric entry angle) also depends on height rate of change of atmospheric density, as characterized by density 
scale height. Characteristics of hydrostatics and the gas law equation mean that relatively small systematic differences in 
temperature-versus-pressure profiles can integrate at high altitudes to very large differences in density-versus-altitude profiles. 
Thus a given periapsis density required to accomplish successful aerocapture can occur at substantially different altitudes 
(approx. 150 - 300 km) on the various outer planets, and significantly different density scale heights (approx. 20 - 50 km) can 
occur at these periapsis altitudes. This paper will illustrate these effects and discuss implications for improvements in 
atmospheric measurements to yield significant impact on design of aerocapture systems for future missions to Titan and the 
outer planets. Relatively small- scale atmospheric perturbations, such as gravity waves, tides, and other atmospheric variations 
can also have significant effect on design details for aerocapture guidance and control systems. This paper will also discuss 
benefits that would result from improved understanding of Titan and outer planetary atmospheric perturbation characteristics. 
Details of recent engineering-level atmospheric models for Titan and Neptune will be presented, and effects of present and 
future levels of atmospheric uncertainty and variability characteristics will be examined. 
Author 
Atmospheric Physics; Atmospheric Models; Aerocapture; Planetary Atmospheres; Atmospheric Density; Annual Variations 

20040062499 Lunar and Planetary Inst., Houston, TX, USA 

Lunar and Planetary Science XXXV: Missions and Instruments: Hopes and Hope Fulfilled 

2004; In English; Lunar and Planetary Science XXXV, 15-19 Mar. 2004, Houston, TX, USA 

Contract(s)/Grant(s): NCC5-679 

Report No.(s): LPI-Contrib-1197; Copyright; Avail: CASI; C01, CD-ROM 

The titles in this section include: 1) Mars Global Surveyor Mars Orbiter Camera in the Extended Mission: The MOC 
Toolkit; 2) Mars Odyssey THEMIS-VIS Calibration; 3) Early Science Operations and Results from the ESA Mars Express 
Mission: Focus on Imaging and Spectral Mapping; 4) The Mars Express/NASA Project at JPL; 5) Beagle 2: Mission to Mars 
- Current Status; 6) The Beagle 2 Microscope; 7) Mars Environmental Chamber for Dynamic Dust Deposition and Statics 
Analysis; 8) Locating Targets for CRISM Based on Surface Morphology and Interpretation of THEMIS Data; 9) The Phoenix 
Mission to Mars; 10) First Studies of Possible Landing Sites for the Phoenix Mars Scout Mission Using the BMST; 11) The 
2009 Mars Telecommunications Orbiter; 12) The Aurora Exploration Program - The ExoMars Mission; 13) Electron-induced 
Luminescence and X-Ray Spectrometer (ELXS) System Development; 14) Remote-Raman and Micro-Raman Studies of Solid 
C02, CH4, Gas Hydrates and Ice; 15) The Compact Microimaging Spectrometer (CMIS): A New Tool for In-Situ Planetary 
Science; 16) Preliminary Results of a New Type of Surface Property Measurement Ideal for a Future Mars Rover Mission; 
17) Electrodynamic Dust Shield for Solar Panels on Mars; 18) Sensor Web for Spatio-Temporal Monitoring of a Hydrological 
Environment; 19) Field Testing of an In-Situ Neutron Spectrometer for Planetary Exploration: First Results; 20) A Miniature 
Solid-State Spectrometer for Space Applications - Field Tests; 21) Application of Laser Induced Breakdown Spectroscopy 
(LIBS) to Mars Polar Exploration: LIBS Analysis of Water Ice and Water Ice/Soil Mixtures; 22) LIBS Analysis of Geological 
Samples at Low Pressures: Application to Mars, the Moon, and Asteroids; 23) In-Situ 1-D and 2-D Mapping of Soil Core and 
Rock Samples Using the LIBS Long Spark; 24) Rocks Analysis at Stand Off Distance by LIBS in Martian Conditions; 25) 
Evaluation of a Compact Spectrograph/Detection System for a LIBS Instrument for In-Situ and Stand-Off Detection; 26) 



Analysis of Organic Compounds in Mars Analog Samples; 27) Report of the Organic Contamination Science Steering Group; 
28) The Water- Wheel IR (WIR) - A Contact Survey Experiment for Water and Carbonates on Mars; 29) Mid-IR Fiber Optic 
Probe for In Situ Water Detection and Characterization; 30) Effects of Subsurface Sampling & Processing on Martian Simulant 
Containing Varying Quantities of Water; 31) The Subsurface Ice Probe (SIPR): A Low-Power Thermal Probe for the Martian 
Polar Layered Deposits; 32) Deploying Ground Penetrating Radar in Planetary Analog Sites to Evaluate Potential Instrument 
Capabilities on Future Mars Missions; 33) Evaluation of Rock Powdering Methods to Obtain Fine-grained Samples for 
CHEMIN, a Combined XRD/XRF Instrument; 34) Novel Sample-handling Approach for XRD Analysis with Minimal Sample 
Preparation; 35) A New Celestial Navigation Method for Mars Landers; 36) Mars Mineral Spectroscopy Web Site: A Resource 
for Remote Planetary Spectroscopy. 
CASI 
Spacecraft Instruments; Planetology; Mars Missions 

20040039671 

Multibody Parachute Flight Simulations for Planetary Entry Trajectories Using 'Equilibrium Points' 

Raiszadeh, Ben; Advances in the Astronautical Sciences; 2003; ISSN 0065-3438; Volume 114, Issue SUPPL., p. 903-914; In 
English; Copyright; Avail: Other Sources 

A method has been developed to reduce numerical stiffness and computer CPU requirements of high fidelity multibody 
flight simulations involving parachutes for planetary entry trajectories. Typical parachute entry configurations consist of entry 
bodies suspended from a parachute, connected by flexible lines. To accurately calculate line forces and moments, the 
simulations need to keep track of the point where the flexible lines meet (confluence point). In previous multibody parachute 
flight simulations, the confluence point has been modeled as a point mass. Using a point mass for the confluence point tends 
to make the simulation numerically stiff, because its mass is typically much less that than the main rigid body masses. One 
solution for stiff differential equations is to use a very small integration time step. However, this results in large computer CPU 
requirements. In the method described in the paper, the need for using a mass as the confluence point has been eliminated. 
Instead, the confluence point is modeled using an 'equilibrium point' . This point is calculated at every integration step as the 
point at which sum of all line forces is zero (static equilibrium). The use of this 'equilibrium point' has the advantage of both 
reducing the numerical stiffness of the simulations, and eliminating the dynamical equations associated with vibration of a 
lumped mass on a high-tension string. 
EI 
Computers; Parachutes; Spacecraft; Stiffness; Trajectories 

20040039371 

Approach navigation for the 2009 Mars large lander 

Burkhart, P. Daniel; Advances in the Astronautical Sciences; 2003; ISSN 0065-3438; Volume 114, Issue SUPPL., p. 
2181-2196; In English; Copyright; Avail: Other Sources 

The current Mars exploration plan envisions the launch of a large lander in the 2009 launch opportunity with a soft landing 
on Mars in the fall of 2010. The goal is to achieve a landed surface position within 10km of the target landing site. Current 
entry descent and landing (EDL) analysis shows that the largest contributor to the landed position error is uncertainty of the 
initial conditions, which are supplied by the ground-based navigation process. The focus of this paper is the performance of 
the approach navigation process using combinations of Deep Space Network (DSN) Doppler, ranging and delta differential 
one-way range (delta DOR) measurements along with optical navigation data collected by the spacecraft. Results for several 
combinations of data types will be included. 
EI 
Data Acquisition; Navigation; Planetary Landing; Roving Vehicles; Trajectories 

20040039331 

Optical landmark detection for spacecraft navigation 

Cheng, Yang; Johnson, Andrew E.; Matthies, Larry H.; Olson, Clark F; Advances in the Astronautical Sciences; 2003; ISSN 
0065-3438; Volume 114, Issue SUPPL., p. 1767-1785; In English; Copyright; Avail: Other Sources 

Optical landmark navigation using craters on the surface of a central body was first used operationally by the Near Earth 
Asteroid Rendezvous (NEAR) mission. It has proven to be a powerful data type for determining spacecraft orbits above the 
target for close flybys and low altitude orbiting. Tracking individual landmarks, which are small craters, enables orbit 
determination accuracies on the order of the camera resolution or several meters. This exceeds the accuracy that can be 



obtained from radiometric data alone. Currently, most of optical landmark navigation operations, such as crater detection, 
tracking, and matching etc, are done manually, which is extremely time consuming, tedious and sometime unmanageable. 
Because of the lengthily operation time and the deep-space communication delay, manual operation cannot meet the 
requirements of rapid and precise spacecraft maneuvers such as close orbiting, fast fiybys and landing. Automating this 
operation can greatly improve navigation accuracy and efficiency and ultimately lead to an on-board autonomous navigation 
capability. In this paper, a new crater detection algorithm is suggested. Experimental studies show that this new algorithm can 
achieve sub-pixel accuracy in position, its detection rate is better than 90% and its false alarm rate is less than 5%. These good 
characteristics indicate that it is an ideal crater detection algorithm for spacecraft optical navigation. 
EI 
Cameras; Navigation; Resolution; Spacecraft Propulsion; Tracking (Position) 



20040039275 

Daily repeat-groundtrack Mars orbits 

Noreen, Gary; Kerridge, Stuart; Diehl, Roger; Neelon, Joseph; Ely, Todd; Turner, Andrew E.; Advances in the Astronautical 
Sciences; 2003; ISSN 0065-3438; Volume 114, Issue SUPPL., p. 1143-1155; In English; Copyright; Avail: Other Sources 

This paper derives orbits at Mars with groundtracks that repeat at the same times every solar day (sol). A relay orbiter 
in such an orbit would pass over in-situ probes at the same times every sol, ensuring consistent coverage and simplifying 
mission design and operations. 42 orbits in five classes are characterized: 14 circular equatorial prograde orbits 14 circular 
equatorial retrograde orbits 1 1 circular sun synchronous orbits 2 eccentric equatorial orbits 1 eccentric critically inclined orbit 
The paper reports on the performance of a relay orbiter in some of the orbits. 
EI 
Aerospace Sciences; Communication Satellites; Ground Tracks; Orbits; Planetary Landing; Planets 



20040038205 

Entry trajectory and atmosphere reconstruction methodologies for the mars exploration rover mission 

Desai, Prasun N.; Blanchard, Robert C; Powell, Richard W.; European Space Agency, (Special Publication) ESA SP; February 
2004; ISSN 0379-6566, Issue no. 544, p. 213-220; In English; International Workshop: Planetary Probe Atmospheric Entry 
and Descent Trajectory Analysis and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources 

The Mars Exploration Rover (MER) mission will land two landers on the surface of Mars, arriving in January 2004. Both 
landers will deliver the rovers to the surface by decelerating with the aid of an aeroshell, a supersonic parachute, retro-rockets, 
and air bags for safely landing on the surface. The reconstruction of the MER descent trajectory and atmosphere profile will 
be performed for all the phases from hypersonic flight through landing. A description of multiple methodologies for the flight 
reconstruction is presented from simple parameter identification methods through a statistical Kalman filter approach. 
EI 
Air Bag Restraint Devices; Kalman Filters; Parachutes; Planetary Landing; Trajectories 



20040038193 

Entry descent, and landing scenario for the Mars exploration Rover mission 

Desai, Prasun N; Lee, Wayne J.; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, 
Issue no. 544, p. 31-36; In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory 
Analysis and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources 

In January 2004, the Mars Exploration Rover (MER) mission will land two landers on the surface of Mars. Both landers 
will deliver a rover to the surface using an entry, descent, and landing (EDL) scenario based on Mars Pathfinder heritage. 
However, the entry conditions and environments are different from that of Mars Pathfinder. Unique challenges are present due 
to the entry differences of a heavier entry mass, less dense atmosphere, and higher surface landing site altitude. These 
differences result in a higher terminal velocity and less time for performing all the EDL events as compared to Mars Pathfinder. 
As a result of these differences, modifications are made to the MER EDL systems to safely deliver the rovers to the surface 
of Mars. 
EI 
Aerospace Sciences; Planetary Landing; Planets; Roving Vehicles; Topography 



20040038111 

Thermal protection system technology and facility needs for demanding future planetary missions 

Laub, B.; Venkatapathy, E.; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, Issue 
no. 544, p. 239-247; In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory Analysis 
and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources 

NASA has successfully launched numerous science missions to inner and outer planets in our solar system of which the 
most challenging were to Venus and Jupiter and the knowledge gained from those missions have been invaluable yet 
incomplete. Future missions will be built on what we have learned from the past missions but they will be more demanding 
from both the science as well as the mission design and engineering perspectives. The Solar System Exploration Decadal 
Survey (SSEDS) produced for NASA by the National Research Council identified a broad range of science objectives many 
of which can only be satisfied with atmospheric entry probes. The SSEDS recommended new probe/lander missions to both 
Venus and Jupiter. The Pioneer- Venus probe mission was launched in August 1978 and four probes successfully entered the 
Venusian atmosphere in December 1978. The Galileo mission was launched in October 1989 and one probe successfully 
entered the Jovian atmosphere in December 1995. The thermal protection system requirements for these two missions were 
unlike any other planetary probes and required fully dense carbon phenolic for the forebody heat shield. Developing thermal 
protection systems to accomplish future missions outlined in the Decadal Survey presents a technology challenge since they 
will be more demanding than these past missions. Unlike Galileo, carbon phenolic may not be an adequate TPS for a future 
Jupiter multiprobe mission since non-equatorial probes will enter at significantly higher velocity than the Galileo equatorial 
probe and the entry heating scales approximately with the cube of the entry velocity. At such heating rates the TPS mass 
fraction for a carbon phenolic heat shield would be prohibitive. A new, robust and efficient TPS is required for such probes. 
The Giant Planet Facility (GPF), developed and employed during the development of the TPS for the Galileo probe was 
dismantled after completion of the program. Furthermore, flight data from the Galileo probe suggested that the complex 
physics associated with the interaction between massive ablation and a severe shock layer radiation environment is not well 
understood or modeled. The lack of adequate ground test facilities to support the development and qualification of new TPS 
materials adds additional complexities. The requirements for materials development, ground testing and sophisticated 
modeling to enable these challenging missions are the focus of this paper. 
EI 
Aerospace Sciences; Heat Shielding; Planetary Landing; Space Probes; Vaporizing 

20040038093 

Ultra-stable oscillators for planetary entry probes 

Asmar, S. W.; Atkinson, D. H.; Bird, M. K.; Wood, G. E.; European Space Agency, (Special Publication) ESA SP; February 
2004; ISSN 0379-6566, Issue no. 544, p. 131-134; In English; International Workshop: Planetary Probe Atmospheric Entry 
and Descent Trajectory Analysis and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources 

Ultra-stable oscillators on-board planetary missions were developed for Radio Science instrumentation, functioning as 
frequency references for the one-way downlink during atmospheric occultations. They have also been flown on planetary entry 
probes including the Jupiter entry probe, carried by Galileo, and the Huygens Titan entry probe, carried by Cassini, for 
performing Doppler Wind Experiments. The Jupiter and Titan probes utilized different oscillators, quartz and rubidium, 
respectively. This paper presents the development of ultra-stable oscillators on deep space missions and discusses the tradeoffs 
encountered when selecting oscillators for planetary entry probes, including factors such as duration of the experiment, the 
available warm-up time and the Allan deviation and phase noise requirements. 
EI 
Mechanical Oscillators; Quartz; Rubidium; Transponders 

20040038075 

Pioneer Venus and Galileo entry probe heritage 

Bienstock, Bernard J.; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, Issue no. 
544, p. 37-45; In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and 
Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources 

Beginning in the late 1960s, NASA began planning for its first program to explore Venus. Although planetary entry probes 
had been flown to Venus by the Soviets beginning in 1967, NASA had not previously flown this type of mission, The Space 
and Communications Group of Hughes Aircraft Company, now owned by Boeing and called Boeing Satellite Systems, worked 
with NASA to perform initial studies that culminated with a contract for the Pioneer Venus program in early 1974, Pioneer 
Venus was an ambitious program that included four planetary entry probes, transported to Venus by a Multiprobe Bus, and 



a Venus Orbiter. This paper focuses on the engineering aspects of the probes and the challenges overcome in accommodating 
the various scientific instruments. The second NASA planetary entry program was the Galileo Mission that began with initial 
studies in the early 1970s. This mission to Jupiter included both an Orbiter and a Probe. Although the Galileo Probe planetary 
entry program was begun as the Pioneer Venus probes were heading towards Venus, there were significant engineering 
differences between the Pioneer Venus probe designs and the Galileo Probe. These differences, dictated by a number of factors, 
are discussed. The paper concludes with a summary of lessons learned by Boeing and NASA in designing, manufacturing and 
ultimately flying the Venus and Jupiter planetary entry probes. 
EI 
Aerospace Sciences; Galileo Spacecraft; Planets; Pressure Vessels; Space Probes; Venus (Planet) 

20040038071 

Summary of the Boulder Entry Probe Workshop April 21-22, 2003, Boulder, Colorado, USA 

Young, Richard E.; Atkinson, David; Atreya, Sushil; Banfield, Donald; Beebe, Reta; Bolton, Scott; Briggs, Geoffrey; Crisp, 
David; Cutts, James; Drake, Michael; Esposito, Larry; Galal, Kenneth; Hubbard, William; Hunten, Donald; Ingersoll, Andrew; 
et al., T; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, Issue no. 544, p. 13-20; 
In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and Science, Oct. 
6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources 

The Solar System Exploration Decadal Survey (SSEDS) identified several high priority Solar System Key Science 
Questions that should be addressed by entry probes/landers, or that should be addressed simultaneously by both probes/landers 
and remote sensing types of observations. These Key Science Questions are directly relevant to Goals and Objectives of the 
NASA Strategic Plan and Office of Space Science Strategic Plan. In this report we define entry probes/landers as spacecraft 
that sample in-situ a planetary atmosphere, and planetary surface if there is one. The Entry Probe Workshop grew out of a 
community concern that if entry probes/landers were to be a viable mission option for addressing the overarching questions 
identified in the SSEDS in the coming 10-15 years, significant resources must be applied to key technology areas immediately. 
The major science objectives requiring entry probes and the key technology development areas for probes are described. 
EI 
Aerospace Sciences; Meteorites; Periodic Variations; Planetary Landing; Planets; Space Probes 

20040037789 NASA Langley Research Center, Hampton, VA, USA 
Angle-of-Attack-Modulated Terminal Point Control for Neptune Aerocapture 

Queen, Eric M.; [2004]; In English, 8-12 Feb. 2004, Maui, HI, USA 

Contract(s)/Grant(s): 800-90-50 

Report No.(s): AAS-04-129; Copyright; Avail: CASI; A02, Hardcopy 

An aerocapture guidance algorithm based on a calculus of variations approach is developed, using angle of attack as the 
primary control variable. Bank angle is used as a secondary control to alleviate angle of attack extremes and to control 
inclination. The guidance equations are derived in detail. The controller has very small onboard computational requirements 
and is robust to atmospheric and aerodynamic dispersions. The algorithm is applied to aerocapture at Neptune. Three versions 
of the controller are considered with varying angle of attack authority. The three versions of the controller are evaluated using 
Monte Carlo simulations with expected dispersions. 
Author 
Algorithms; Aerocapture; Angle of Attack; Neptune (Planet); Control Theory; Terminal Guidance 

20040037788 NASA Langley Research Center, Hampton, VA, USA 

Mars Exploration Rover Terminal Descent Mission Modeling and Simulation 

Raiszadeh, Behzad; Queen, Eric M.; February 2004; In English, 8-12 Feb. 2004, Maui, HI, USA 

Contract(s)/Grant(s): 759-30-00 

Report No.(s): AAS-04-271; No Copyright; Avail: CASI; A03, Hardcopy 

Because of NASA's added reliance on simulation for successful interplanetary missions, the MER mission has developed 
a detailed EDL trajectory modeling and simulation. This paper summarizes how the MER EDL sequence of events are 
modeled, verification of the methods used, and the inputs. This simulation is built upon a multibody parachute trajectory 
simulation tool that has been developed in POST II that accurately simulates the trajectory of multiple vehicles in flight with 
interacting forces. In this model the parachute and the suspended bodies are treated as 6 Degree-of-Freedom (6 DOF) bodies. 
The terminal descent phase of the mission consists of several Entry, Descent, Landing (EDL) events, such as parachute 



deployment, heatshield separation, deployment of the lander from the backshell, deployment of the airbags, RAD firings, TIRS 

firings, etc. For an accurate, reliable simulation these events need to be modeled seamlessly and robustly so that the 

simulations will remain numerically stable during Monte-Carlo simulations. This paper also summarizes how the events have 

been modeled, the numerical issues, and modeling challenges. 

Author 

Mars Exploration; Mars Roving Vehicles; Descent; Space Missions; Mathematical Models; Trajectory Analysis; 

Computerized Simulation 

20040024535 

Planning for a Mars in situ sample preparation and distribution (SPAD) system 

Beaty, D. W.; Miller, S.; Zimmerman, W.; Bada, J.; Conrad, P.; Dupuis, E.; Huntsberger, T.; Ivlev, R.; Kim, S. S.; Lee, B. G.; 
Lindstrom, D.; Lorenzoni, L.; Mahaffy, P.; McNamara, K.; Papanastassiou, D.; et al., T; Planetary and Space Science; 
January/March 2004; ISSN 0032-0633; Volume 52, Issue no. 1-3, p. 55-66; In English; Copyright; Avail: Other Sources 

For Mars in situ landed missions, it has become increasingly apparent that significant value may be provided by a shared 
system that we call a Sample Preparation and Distribution (SPAD) System. A study was conducted to identify the issues and 
feasibility of such a system for these missions that would provide common functions for: receiving a variety of sample types 
from multiple sample acquisition systems; conducting preliminary characterization of these samples with non-destructive 
science instruments and making decisions about what should happen to the samples; performing a variety of sample 
preparation functions; and, finally, directing the prepared samples to additional science instruments for further analysis. 
Scientific constraints on the functionality of the system were identified, such as triage, contamination management, and 
various sample preparation steps, e.g., comminution, splitting, rock surfacing, and sieving. Some simplifying strategies were 
recommended and an overall science flow was developed. Engineering functional requirements were also investigated and 
example architectures developed. Preliminary conclusions are that shared SPAD facility systems could indeed add value to 
future Mars in situ landed missions if they are designed to respond to the particular requirements and constraints of those 
missions, that such a system appears feasible for consideration, and that certain standards should be developed for key SPAD 
interfaces, (copyright) 2003 Elsevier Ltd. All rights reserved. 
EI 
Aerospace Sciences; In Situ Measurement; Planetary Landing; Spacecraft 

20040024261 

Blended control, predictor-corrector guidance algorithm: An enabling technology for Mars aerocapture 

Jits, Roman Y.; Walberg, Gerald D.; Acta Astronautica; March 2004; ISSN 0094-5765; Volume 54, Issue no. 6, p. 385-398; 
In English; Copyright; Avail: Other Sources 

A guidance scheme designed for coping with significant dispersion in the vehicle's state and atmospheric conditions is 
presented. In order to expand the flyable aerocapture envelope, control of the vehicle is realized through bank angle and 
angle-of-attack modulation. Thus, blended control of the vehicle is achieved, where the lateral and vertical motions of the 
vehicle are decoupled. The overall implementation approach is described, together with the guidance algorithm macrologic 
and structure. Results of guidance algorithm tests in the presence of various single and multiple off-nominal conditions are 
presented and discussed, (copyright) 2003 Published by Elsevier Ltd. 
EI 
Aerospace Sciences; Astrophysics; Atmospheric Chemistry; Interplanetary Spacecraft; Planets; Predictor-Corrector Methods 

20040012726 NASA Marshall Space Flight Center, Huntsville, AL, USA 

SEP Mission to Titan NEXT Aerocapture In-Space Propulsion (Quicktime Movie) 

Baggett, Randy; TECH ISP: Next Generation Ion; January 2004; In English; No Copyright; Avail: CASI; A01, Hardcopy 

The ion thruster is one of the most promising solar electric propulsion (SEP) technologies to support future Outer Planet 
missions (place provided link below here) for NASA's Office of Space Science. Typically, ion thrusters are used in high Isp- 
low thrust applications that require long lifetimes, as well as, higher efficiency over state-of-the-art chemical propulsion 
systems. Today, the standard for ion thrusters is the SEP Technology Application Readiness (NSTAR) thruster. Jet Propulsion 
Laboratory's (JPL's) extended life test (ELT) of the DS 1 flight spare NSTAR thruster began in October 1998. This test 
successfully demonstrated lifetime of the NSTAR flight spare thruster, which will provide a solid basis for selection of ion 
thrusters for future Code S missions. The NSTAR ELT was concluded on June 30,2003 after 30,352 hours. The purpose of 
the Next Generation Ion (NGI) activities is to advance Ion propulsion system technologies through the development of 

10 



NASA's Evolutionary Xenon Thruster (NEXT). The goal of NEXT is to more than double the power capability and lifetime 
throughput (the total amount of propellant which can be processed) while increasing the Isp by 30% and the thrust by 120%. 
Derived from text 
Ion Propulsion; Solar Electric Propulsion 

20030111896 Naval Postgraduate School, Monterey, CA 
Optimization of Low Thrust Trajectories With Terminal Aerocapture 

Josselyn, Scott B.; Jun. 2003; In English; Original contains color illustrations 
Report No.(s): AD-A417512; No Copyright; Avail: CASI; A08, Hardcopy 

This thesis explores using a direct pseudospectral method for the solution of optical control problems with mixed 
dynamics. An easy to use MATLAB optimization package known as DIDO is used to obtain the solutions. The modeling of 
both low thrust interplanetary trajectories as well as aerocapture trajectories is detailed and the solutions for low thrust 
minimum time and minimum fuel trajectories are explored with particular emphasis on verification of the optimality of the 
obtained solution. Optimal aerocapture trajectories are solved for rotating atmospheres over a range of arrival V- infinities. 
Solutions are obtained using various performance indexes including minimum fuel, minimum heat load, and minimum total 
aerocapture mass. Finally, the problem formulation and solutions for the mixed dynamic problem of low thrust trajectories 
with a terminal aerocapture maneuver is addressed yielding new trajectories maximizing the total scientific mass at arrival. 
This thesis explores using a direct pseudospectral method for the solution of optimal control problems with mixed dynamics. 
An easy to use MATLAB optimization package known as DIDO is used to obtain the solutions. The modeling of both low 
thrust interplanetary trajectories as well as aerocapture trajectories is detailed and the solutions for low thrust minimum time 
and minimum fuel trajectories are explored with particular emphasis on verification of the optimality of the obtained solution. 
Optimal aerocapture trajectories are solved for rotating atmospheres over a range of arrival V-infmities. Solutions are obtained 
using various performance indexes including minimum fuel, minimum heat load, and minimum total aerocapture mass. 
Finally, the problem formulation and solutions for the mixed dynamic problem of low thrust trajectories with a terminal 
aerocapture maneuver is addressed yielding new trajectories maximizing the total scientific mass at arrival. 
DTIC 
Interplanetary Trajectories; Interorbital Trajectories; Trajectory Optimization; Optimal Control 

20030107097 Air Force Inst, of Tech., Wright-Patterson AFB, OH 
Aerocapture Guidance Methods for High Energy Trajectories 

Dicarlo, Jennifer L.; May 23, 2003; In English 

Report No.(s): AD-A416545; AFIT-CI02-1191; No Copyright; Avail: CASI; A07, Hardcopy 

This thesis investigates enhancements of an existing numerical predictor-corrector aerocapture guidance algorithm 
(PredGuid). The study includes implementation of an energy management phase prior to targeting with a generic method of 
transition and replacement of heuristic features with more generic features. The vehicle response during energy management 
was modeled as a second-order spring/mass/damper system. Phase change occurred when two conditions were met: First, the 
vehicle could fly a constant bank angle of 1100 for the remainder of the trajectory and have the resulting apogee below or 
within a given tolerance above the target apogee. Second, the predicted final energy indicated that the vehicle would be on 
an elliptical, not hyperbolic, trajectory. So as to incorporate generic features, modeling of a separate lift down phase was 
replaced by using a lift-down condition to determine phase change and biasing to the same lift- down condition during 
targeting. Also, use of a heuristic sensitivity to calculate the first corrected bank angle was replaced by a simple smart 
guessing' algorithm. Finally, heuristic lateral corridor boundaries were replaced by boundaries based on percentage of forward 
velocity. 
DTIC 
Trajectories; Hyperbolic Trajectories; Predictor-Corrector Methods; Algorithms 

20030106653 NASA Marshall Space Flight Center, Huntsville, AL, USA 
Aerocapture Technology Project Overview 

James, Bonnie; Munk, Michelle; Moon, Steve; July 20, 2003; In English 

Report No.(s): AIAA Paper 2003-4654; No Copyright; Avail: CASI; A01, Hardcopy 

Aerocapture technology development is one of the highest priority investments for the NASA In-Space Propulsion 
Program (ISP). The ISP is managed by the NASA Headquarters Office of Space Science, and implemented by the Marshall 
Space Flight Center in Huntsville, Alabama. The objective of the ISP Aerocapture Technology Project (ATP) is to develop 

11 



technologies that can enable and/or benefit NASA science missions by significantly reducing cost, mass, and trip times. To 

accomplish this objective, the ATP identifies and prioritizes the most promising technologies using systems analysis, 

technology advancement and peer review, coupled with NASA Headquarters Office of Space Science target requirements. 

Efforts are focused on developing mid-Technology Readiness Level (TRL) technologies to systems-level spaceflight 

validation. 

Author 

Aerocapture; Systems Analysis; Spacecraft Propulsion 

20030106138 Ball Aerospace and Technologies Corp., Boulder, CO, USA 
Trailing Ballute Aerocapture: Concept and Feasibility Assessment 

Miller, Kevin L.; Gulick, Doug; Lewis, Jake; Trochman, Bill; Stein, Jim; Lyons, Daniel T; Wilmoth, Richard G.; July 21, 
2003; In English; AIAA Joint Propulsion Conference and Exhibit 2003, 20-23 Jul. 2003, Huntsville, AL, USA 
Contract(s)/Grant(s): NAS8-02130; JPL-1205966 
Report No.(s): AIAA Paper 2003-4655; Copyright; Avail: CASI; A03, Hardcopy 

Trailing Ballute Aerocapture offers the potential to obtain orbit insertion around a planetary body at a fraction of the mass 
of traditional methods. This allows for lower costs for launch, faster flight times and additional mass available for science 
payloads. The technique involves an inflated ballute (balloon-parachute) that provides aerodynamic drag area for use in the 
atmosphere of a planetary body to provide for orbit insertion in a relatively benign heating environment. To account for 
atmospheric, navigation and other uncertainties, the ballute is oversized and detached once the desired velocity change (Delta 
V) has been achieved. Analysis and trades have been performed for the purpose of assessing the feasibility of the technique 
including aerophysics, material assessments, inflation system and deployment sequence and dynamics, configuration trades, 
ballute separation and trajectory analysis. Outlined is the technology development required for advancing the technique to a 
level that would allow it to be viable for use in space exploration missions. 
Author 
Ballutes; Aerocapture; Aerodynamic Drag; Feasibility Analysis 

20030105420 

Mars reconnaissance orbiter design approach for high-resolution surface imaging 

Lee, S. W.; Skulsky, E. D.; Chapel, J.; Cwynar, D.; Gehling, R.; Delamere, A.; Advances in the Astronautical Sciences; 2003; 
ISSN 0065-3438; Volume 113, p. 509-528; In English; Guidance and Control 2003: Advances in the Astronautical Sciences, 
Feb. 5-9, 2003, Breckenridge, CO, USA; Copyright; Avail: Other Sources 

The Mars Reconnaissance Orbiter (MRO) will explore Mars equipped with a suite of six scientific instruments and two 
engineering experiments, and supporting two additional facility investigations. One of the objectives of the MRO mission is 
to use the High-Resolution Imaging Science Experiment (HiRISE) to provide 30 cm/pixel images of future Mars landing sites. 
To achieve such detail, MRO must meet some very challenging target-relative pointing and pointing stability requirements. 
A combination of analysis, operational constraints, and spacecraft design modifications were utilized to ensure that the 
necessary pointing requirements will be met. 
EI 
High Resolution; Imaging Techniques; Orbits; Reconnaissance Aircraft 

20030091868 

Pitch control during autonomous aerobraking for near-term Mars exploration 

Johnson, Wyatt R.; Longuski, James M.; Lyons, Daniel T; Journal of Spacecraft and Rockets; May/June 2003; ISSN 
0022-4650; Volume 40, Issue no. 3, p. 371-379; In English; Copyright; Avail: Other Sources 

Conventional aerobraking requires propellant to dump the spacecraft's angular momentum and to maintain attitude 
control during the atmospheric fly through. We consider how reaction wheels can be used to control the spacecraft's pitch 
during each atmospheric flythrough and to reduce angular momentum simultaneously. Control laws are developed for 
minimum onboard instrumentation (where the only state information are the angular rates of the spacecraft and the reaction 
wheels) to compensate for large variations in entry time and atmospheric density. Simulations indicate that pitch attitude and 
angular momentum can be controlled with reaction wheels alone, thus saving precious propellant while significantly 
increasing the timing margin for sequencing. 
EI 
Aerodynamics; Computerized Simulation; Drag; Spacecraft 

12 



20030080878 

AIMS: Acousto-optic imaging spectrometer for spectral mapping of solid surfaces 

Glenar, David A.; Blaney, Diana L.; Hillman, John J.; Acta Astronautica; January/March 2003; ISSN 0094-5765; Volume 52, 
Issue no. 2-6, p. 389-396; In English; Copyright; Avail: Other Sources 

A compact, two-channel acousto-optic tunable filter (AOTF) camera is being built at GSFC as a candidate payload 
instrument for future Mars landers or small-body rendezvous missions. This effort is supported by the NASA Mars Instrument 
Development Program (MIDP), Office of Space Science Advanced Technologies and Mission Studies. Acousto-optic Imaging 
Spectrometer (AIMS) is electronically programmable and provides arbitrary spatial and spectral selection from 0.48 to 2.4 mu 
m. The geometric throughput of AOTF' s are well matched to the requirements for lander mounted cameras since (I) they can 
be made very compact, (II) 'slow' (f/14-f/18) optics required for large depth-of-field fall well within the angular aperture limit 
of AOTF' s, and (III) they operate at low ambient temperatures. A breadboard of the AIMS short-wavelength channel is now 
being used for spectral imaging of high-interest Mars analog materials (iron oxides, carbonates, sulfates and sedimentary 
basalts) as part of the initial instrument validation exercises, (copyright) 2002 Published by Elsevier Science Ltd. 
EI 
Acousto-Optics; Aerospace Sciences; Cameras; Imaging Techniques; Planetary Landing 

20030080863 
Europa Lander 

Gershman, Robert; Nilsen, Erik; Oberto, Robert; Acta Astronautica; January/March 2003; ISSN 0094-5765; Volume 52, Issue 
no. 2-6, p. 253-258; In English; Copyright; Avail: Other Sources 

A Europa Lander mission has been assigned high priority for the post-2005 time frame in NASA's Space Science 
Enterprise Strategic Plan. Europa is one of the most scientifically interesting objects in the solar system because of the strong 
possibility that a liquid water ocean exists underneath its ice-covered surface. The primary scientific goals of the proposed 
Europa Lander mission are to characterize the surface material from a recent outflow and look for evidence of pre-biotic and 
possibly biotic chemistry. The baseline mission concept involves landing a single spacecraft on the surface of Europa with the 
capability to acquire samples of material, perform detailed chemical analysis of the samples, and transmit the results to Earth. 
This paper provides a discussion of the benefits and status of the key spacecraft and instrument technologies needed to 
accomplish the science objectives. Also described are variations on the baseline concept including the addition of small 
auxiliary probes and an experimental ice penetration probe, (copyright) 2002 Elsevier Science Ltd. All rights reserved. 
EI 
Aerospace Sciences; Planetary Landing; Solar System; Spacecraft 

20030066242 NASA Marshall Space Flight Center, Huntsville, AL, USA 
Engineering-Level Model Atmospheres for Titan & Neptune 

Justus, C. G.; Johnson, D. L.; July 20, 2003; In English, 20-23 Jul. 2003, Huntsville, AL, USA 
Contract(s)/Grant(s): NAS8-60000; No Copyright; Avail: CASI; A01, Hardcopy 

Engineering-level atmospheric models for Titan and Neptune have been developed for use in NASA s systems analysis 
studies of aerocapture applications in missions to the outer planets. Analogous to highly successful Global Reference 
Atmospheric Models for Earth (GRAM, Justus et al., 2000) and Mars (Mars-GRAM, Justus and Johnson, 2001, Justus et al., 
2002) the new models are called Titan-GRAM and Neptune-GRAM. Like GRAM and Mars-GRAM, an important feature of 
Titan-GRAM and Neptune-GRAM is their ability to simulate quasi-random perturbations for Monte- Carlo analyses in 
developing guidance, navigation and control algorithms, and for thermal systems design. 
Author 
Aerocapture; Titan; Monte Carlo Method; Neptune (Planet); Atmospheric Models 

20030066102 Rhode Island Univ., Narragansett, RI, USA 
Science and Engineering Potential of an Icy Moon Lander 

DHondt, S. L.; Millerr, J. H.; Forum on Concepts and Approaches for Jupiter Icy Moons Orbiter; 2003, 17; In English; 
Copyright; Abstract Only; Available from CASI only as part of the entire parent document 

We urge consideration of an Icy Moon Lander as part of the Jupiter Icy Moon Orbiter mission. Inclusion of a lander would 
have major advantages. It would allow scientific and engineering objectives to be met that cannot be addressed with an orbiter. 
It would also allow independent tests of surface and subsurface properties inferred from remote observations. It would provide 
invaluable engineering data for the design of a future ice or Ocean penetrator mission. We illustrate these advantages with 

13 



three examples. As the first example, an acoustic profiler imbedded in the surface of an icy moon could be used to identify 
several subsurface properties as a function of depth. Some of these properties, such as the presence and depth (or absence) 
of the water/ice interface and the structure and density of the ice as a function of depth, might be independently inferred by 
instrumentation on an orbiter. Other properties that might be determinable with an acoustic profiler cannot be imaged from 
orbit. These include the shear modulus of the ice (which might be used to distinguish between rigid ice and slushy convecting 
ice), ocean density as a function of depth, the depth of an ocean/bedrock interface, and properties of the bedrock underlying 
the ocean and ice. For the second example, instrumentation on a lander could undertake direct chemical analysis of organic 
and inorganic compounds in the surface ice and atmosphere of an icy moon. Such analyses would directly test models of 
surface compositions and atmospheric composition inferred from remote observations. These analyses would also greatly 
advance human understanding of the chemical habitability of a Jovian icy moon by directly identifying and quantifying 
concentrations of nutrients, energy yielding chemicals, and carbon sources in the surface ice. Thermal studies provide the third 
example. Lander-based thermal measurements on the surface of an icy moon would provide an absolute calibration standard 
for surface temperatures inferred from remote observations. Downhole temperature measurements taken at a single site with 
a shallow penetrator would allow modeling of the subsurface temperature profile and independent estimation of the presence 
and depth (or absence) of the ice/ocean interface. In closing, we wish to emphasize that inclusion of a lander with relatively 
low-weight instrumentation in the JIMO mission would provide a high scientific pay-off. Because the lander instrumentation 
would not penetrate the ice deeply, there would be no risk of directly contaminating any underlying ocean. Such a lander might 
require only modest adaptation of existing technology and consequently might entail relatively low cost. 
Derived from text 
Galilean Satellites; Satellite Surfaces; Surface Properties; Planetary Landing; Measuring Instruments; Space Probes 



20030065170 NASA Langley Research Center, Hampton, VA, USA 

Mars Exploration Rover Six-Degree-Of-Freedom Entry Trajectory Analysis 

Desai, Prasun N.; Schoenenberger, Mark; Cheatwood, F. M.; [2003]; In English, 3-7 Aug. 2003, Big Sky, MT, USA 
Report No.(s): AAS Paper 03-642; Copyright; Avail: CASI; A03, Hardcopy 

The Mars Exploration Rover mission will be the next opportunity for surface exploration of Mars in January 2004. Two 
rovers will be delivered to the surface of Mars using the same entry, descent, and landing scenario that was developed and 
successfully implemented by Mars Pathfinder. This investigation describes the trajectory analysis that was performed for the 
hypersonic portion of the MER entry. In this analysis, a six-degree-of-freedom trajectory simulation of the entry is performed 
to determine the entry characteristics of the capsules. In addition, a Monte Carlo analysis is also performed to statistically 
assess the robustness of the entry design to off-nominal conditions to assure that all entry requirements are satisfied. The results 
show that the attitude at peak heating and parachute deployment are well within entry limits. In addition, the parachute 
deployment dynamics pressure and Mach number are also well within the design requirements. 
Author 

Degrees of Freedom; Trajectory Analysis; Atmospheric Entry; Mars Roving Vehicles; Mars Exploration; NASA Space 
Programs 



20030062242 NASA Marshall Space Flight Center, Huntsville, AL, USA 
NASA Development of Aerocapture Technologies 

James, Bonnie; Munk, Michelle; Moon, Steve; [2003]; In English, 20-22 May 2003, Monterey, CA, USA; Copyright; Avail: 
CASI; A01, Hardcopy 

Aeroassist technology development is a vital part of the NASA ln-Space Propulsion Program (ISP), which is managed 
by the NASA Headquarters Office of Space Science, and implemented by the Marshall Space Flight Center in Huntsville, 
Alabama. Aeroassist is the general term given to various techniques to maneuver a space vehicle within an atmosphere, using 
aerodynamic forces in lieu of propulsive fuel. Within the ISP, the current aeroassist technology development focus is 
aerocapture. The objective of the ISP Aerocapture Technology Project (ATP) is to develop technologies that can enable and/or 
benefit NASA science missions by significantly reducing cost, mass, and/or travel times. To accomplish this objective, the ATP 
identifies and prioritizes the most promising technologies using systems analysis, technology advancement and peer review, 
coupled with NASA Headquarters Office of Space Science target requirements. Plans are focused on developing 
mid-Technology Readiness Level (TRL) technologies to TRL 6 (ready for technology demonstration in space). 
Author 
NASA Space Programs; Aeroassist; Aerodynamic Forces; Systems Analysis; Propulsion; Aerocapture 

14 



20030055137 

Development of a Monte Carlo Mars-gram model for 2001 Mars Odyssey aerobraking simulations 

Dwyer, Alicia M.; Tolson, Robert H.; Munk, Michelle M.; Tartabini, Paul V.; Journal of the Astronautical Sciences; April- June 
2002; ISSN 0021-9142; Volume 50, Issue no. 2, p. 191-211; In English; Copyright 

Atmospheric density data taken during the Mars Global Surveyor aerobraking mission (1997-1999) showed significant 
variability over the altitude range (100-140 km) of interest for aerobraking. This paper presents the method by which Mars 
Global Surveyor data were used to determine the statistical distribution of mean density and the amplitude and phase of 
stationary atmospheric waves as a function of latitude. The combination of mean density and waves produced a good fit to 
the observed data. Using this information, a model was developed to implement the variations into Monte Carlo simulations 
for future missions to Mars, specifically the Mars Odyssey aerobraking mission (October, 2001-January, 2002). An example 
of Monte Carlo results for the Mars Odyssey aerobraking mission is shown. 
EI 
Atmospheric Density; Computerized Simulation; Monte Carlo Method; Planets; Space Flight 

20030055136 

Approaches to autonomous aerobraking at Mars 

Hanna, J. L.; Tolson, R. H.; Journal of the Astronautical Sciences; April-June 2002; ISSN 0021-9142; Volume 50, Issue no. 
2, p. 173-189; In English; Copyright 

Planetary atmospheric aerobraking will most likely be incorporated in every future Mars orbiting mission. Aerobraking 
requires an intensive workload during operations. To provide safe and efficient aerobraking, both navigation and spacecraft 
system teams must be extremely diligent in updating spacecraft sequences and performing periapsis raise or lower maneuvers 
to maintain the required orbital energy reduction without exceeding the design limits of the spacecraft. Automating the process 
with onboard measurements could significantly reduce the operational burden and, in addition, could reduce the potential for 
human error. Two levels of automation are presented and validated using part of the Mars Global Surveyor aerobraking 
sequence and a simulated Mars Odyssey sequence. The simplest method only provides the capability to update the onboard 
sequence. This method uses onboard accelerometer measurements to estimate the change in orbital period during an 
aerobraking pass and thereby estimates the beginning of the next aerobraking sequence. Evaluation of the method utilizing 
MGS accelerometer data showed that the time of the next periapsis can be estimated to within 25% 3 sigma of the change 
in the orbital period due to drag. The second approach provides complete onboard orbit propagation. A low-order gravity 
model is proposed that is sufficient to provide periapsis altitude predictions to within 100-200 meters over three orbits. 
Accelerometer measurements are used as part of the trajectory force model while the spacecraft is in the atmosphere. 
EI 
Accelerometers; Navigation; Planets; Space Flight; Spacecraft 

20030015758 NASA Langley Research Center, Hampton, VA USA 

Wake Closure Characteristics and Afterbody Heating on a Mars Sample Return Orbiter 

Horvath, Thomas J.; Cheatwood, McNeil F.; Wilmoth, Richard G.; Alter, Stephen J.; [2002]; In English, 3-7 Feb. 2002, 
Albuquerque, NM, USA; No Copyright; Avail: CASI; A03, Hardcopy 

Aeroheating wind-tunnel tests were conducted on a 0.028 scale model of an orbiter concept considered for a possible Mars 
sample return mission. The primary experimental objectives were to characterize hypersonic near wake closure and determine 
if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian 
aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the 
NASA Langley 20-Inch Mach 6 Air and CF4 Tunnels for post-normal shock Reynolds numbers (based on forebody diameter) 
ranging from 1,400 to 415,000, angles of attack ranging from -5 to 10 degrees at 0, 3, and 6 degree sideslip, and normal-shock 
density ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was 
inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75 percent of the 
reference forebody stagnation point heating. Comparison of laminar heating prediction to experimental measurement along the 
orbiter afterbody highlight grid alignment challenges associated with numerical simulation of three- dimensional separated 
wake flows. Predicted values of a continuum breakdown parameter revealed significant regions of non-continuum flow 
downstream of the flow separation at the MSRO shoulder and in the region of the reattachment shock on the afterbody. The 
presence of these regions suggest that the Navier-Stokes predictions at the laminar wind-tunnel condition may encounter errors 
in the numerical calculation of the wake shear layer development and impingement due to non-continuum effects. 
Author 

Mars Sample Return Missions; Aerodynamic Heating; Wind Tunnel Tests; Hypersonic Wakes; Impingement; Aerocapture; 
Interplanetary Spacecraft; Flow Characteristics 

15 



20030014800 NASA Langley Research Center, Hampton, VA USA 
Autonomous Aerobraking at Mars 

Hanna, Jill L.; Tolson, Robert; Cianciolo, Alicia Dwyer; Dec, John; [2002]; In English, 22-25 Oct. 2002, Frascati, Italy; 
Original contains color illustrations; Copyright; Avail: CASI; A02, Hardcopy; Distribution as joint owner in the copyright 

Aerobraking has become a proven approach for orbital missions at Mars. A launch of a 1000 kg class spacecraft on a Delta 
class booster saves 90% of the post-MOI fuel otherwise required to circularize the orbit. In 1997, Mars Global Surveyor 
demonstrated the feasibility and Mars 2001 Odyssey completed a nearly trouble free aerobraking phase in January 2002. In 
2006, Mars Reconnaissance Orbiter will also utilize aerobraking. From the flight operations standpoint, however, aerobraking 
is labor intensive and high risk due to the large density variability in the Mars thermosphere. The maximum rate of aerobraking 
is typically limited by the maximum allowable temperature of the solar array which is the primary drag surface. Prior missions 
have used a surrogate variable, usually maximum free stream heat flux, as a basis for performing periapsis altitude corridor 
control maneuvers. This paper provides an adaptive sequential method for operationally relating measured temperatures to 
heat flux profile characteristics and performing maneuvers based directly on measured temperatures and atmospheric 
properties derived from the heat flux profiles. Simulations of autonomous aerobraking are performed using Odyssey mission 
data. 
Author 

Aerobraking; Mars Missions; Spacecraft Maneuvers; Aeromaneuvering; Flight Operations; Computerized Simulation; 
Temperature Profiles; Solar Arrays; Heat Flux 



20030014794 NASA Langley Research Center, Hampton, VA USA 

Multibody Parachute Flight Simulations for Planetary Entry Trajectories Using 'Equilibrium Points' 

Raiszadeh, Ben; [2003]; In English, 9-13 Feb. 2003, Ponce, Puerto Rico; Original contains color illustrations 
Report No.(s): AAS-03-163; No Copyright; Avail: CASI; A03, Hardcopy 

A method has been developed to reduce numerical stiffness and computer CPU requirements of high fidelity multibody 
flight simulations involving parachutes for planetary entry trajectories. Typical parachute entry configurations consist of entry 
bodies suspended from a parachute, connected by flexible lines. To accurately calculate line forces and moments, the 
simulations need to keep track of the point where the flexible lines meet (confluence point). In previous multibody parachute 
flight simulations, the confluence point has been modeled as a point mass. Using a point mass for the confluence point tends 
to make the simulation numerically stiff, because its mass is typically much less that than the main rigid body masses. One 
solution for stiff differential equations is to use a very small integration time step. However, this results in large computer CPU 
requirements. In the method described in the paper, the need for using a mass as the confluence point has been eliminated. 
Instead, the confluence point is modeled using an 'equilibrium point' . This point is calculated at every integration step as the 
point at which sum of all line forces is zero (static equilibrium). The use of this 'equilibrium point' has the advantage of both 
reducing the numerical stiffness of the simulations, and eliminating the dynamical equations associated with vibration of a 
lumped mass on a high-tension string. 
Author 
Atmospheric Entry; Flight Simulation; Parachutes; Trajectories; Differential Equations; Computerized Simulation 



20030014283 NASA Langley Research Center, Hampton, VA USA 
Plume Modeling and Application to Mars 2001 Odyssey Aerobraking 

Chavis, Zachary Q.; Wilmoth, Richard G.; [2002]; In English, 24-26 Jun. 2002, Saint Louis, MO, USA; Original contains 
color illustrations; No Copyright; Avail: CASI; A03, Hardcopy 

A modified source flow model was used to calculate the plume flowfield from a Mars Odyssey thruster during 
aerobraking. The source flow model results compared well with previous detailed CFD results for a Mars Global Surveyor 
thruster. Using an iso-density surface for the Odyssey plume, DSMC simulations were performed to determine the effect the 
plumes have on the Odyssey aerodynamics. A database was then built to incorporate the plume effects into 6-DOF simulations 
over a range of attitudes and densities expected during aerobraking. 6-DOF simulations that included the plume effects showed 
better correlation with flight data than simulations without the plume effects. 
Author 

Computational Fluid Dynamics; Aerobraking; 2001 Mars Odyssey; Computerized Simulation; Rocket Exhaust; Flow 
Distribution; Mathematical Models 

16 



20030006687 NASA Langley Research Center, Hampton, VA USA 

Thermal Analysis and Correlation of the Mars Odyssey Spacecraft's Solar Array During Aerobraking Operations 

Dec, John A.; Gasbarre, Joseph R; George, Benjamin E.; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA; Original 

contains color illustrations 

Report No.(s): AIAA Paper 2002-4536; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose 

rights 

The Mars Odyssey spacecraft made use of multipass aerobraking to gradually reduce its orbit period from a highly 
elliptical insertion orbit to its final science orbit. Aerobraking operations provided an opportunity to apply advanced thermal 
analysis techniques to predict the temperature of the spacecraft's solar array for each drag pass. Odyssey telemetry data was 
used to correlate the thermal model. The thermal analysis was tightly coupled to the flight mechanics, aerodynamics, and 
atmospheric modeling efforts being performed during operations. Specifically, the thermal analysis predictions required a 
calculation of the spacecraft's velocity relative to the atmosphere, a prediction of the atmospheric density, and a prediction 
of the heat transfer coefficients due to aerodynamic heating. Temperature correlations were performed by comparing predicted 
temperatures of the thermocouples to the actual thermocouple readings from the spacecraft. Time histories of the spacecraft 
relative velocity, atmospheric density, and heat transfer coefficients, calculated using flight accelerometer and quaternion data, 
were used to calculate the aerodynamic heating. During aerobraking operations, the correlations were used to continually 
update the thermal model, thus increasing confidence in the predictions. This paper describes the thermal analysis that was 
performed and presents the correlations to the flight data. 
Author 
Thermal Analysis; Correlation; 2001 Mars Odyssey; Solar Arrays; Heat Transfer Coefficients; Aerodynamic Heating 

20030006120 NASA Langley Research Center, Hampton, VA USA 

Control Surface and Afterbody Experimental Aeroheating for a Proposed Mars Smart Lander Aeroshell 

Liechty, Derek S.; Hollis, Brian R.; Edquist, Karl T; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA; Original 

contains color illustrations 

Report No.(s): AIAA Paper 2002-4506; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose 

rights 

Several configurations, having a Viking aeroshell heritage and providing lift-to-drag required for precision landing, have 
been considered for a proposed Mars Smart Lander. An experimental aeroheating investigation of two configurations, one 
having a blended tab and the other a blended shelf control surface, has been conducted at the NASA Langley Research Center 
in the 20-Inch Mach 6 Air Tunnel to assess heating levels on these control surfaces and their effects on afterbody heating. The 
proposed Mars Smart Lander concept is to be attached through its aeroshell to the main spacecraft bus, thereby producing 
cavities in the forebody heat shield upon separation prior to entry into the Martian atmosphere. The effects these cavities will 
have on the heating levels experienced by the control surface and the afterbody were also examined. The effects of Reynolds 
number, angle-of-attack, and cavity location on aeroheating levels and distributions were determined and are presented. At the 
highest angle-of-attack, blended tab heating was increased due to transitional reattachment of the separated shear layer. The 
placement of cavities downstream of the control surface greatly influenced aeroheating levels and distributions. Forebody heat 
shield cavities had no effect on afterbody heating and the presence of control surfaces decreased leeward afterbody heating 
slightly. 
Author 
Control Surfaces; Wind Tunnel Tests; Aeroshells; Aerodynamic Heating; Mars Landing 

20030005808 NASA Langley Research Center, Hampton, VA USA 

The Development and Evaluation of an Operational Aerobraking Strategy for the Mars 2001 Odyssey Orbiter 

Tartabini, Paul V; Munk, Michelle M.; Powell, Richard W.; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA; Original 
contains color illustrations 

Report No.(s): AIAA Paper 2002-4537; No Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government 
purpose rights 

The Mars 2001 Odyssey Orbiter successfully completed the aerobraking phase of its mission on January 11, 2002. This 
paper discusses the support provided by NASA's Langley Research Center to the navigation team at the Jet Propulsion 
Laboratory in the planning and operational support of Mars Odyssey Aerobraking. Specifically, the development of a 
three-degree-of-freedom aerobraking trajectory simulation and its application to pre-flight planning activities as well as 
operations is described. The importance of running the simulation in a Monte Carlo fashion to capture the effects of mission 
and atmospheric uncertainties is demonstrated, and the utility of including predictive logic within the simulation that could 

17 



mimic operational maneuver decision-making is shown. A description is also provided of how the simulation was adapted to 

support flight operations as both a validation and risk reduction tool and as a means of obtaining a statistical basis for 

maneuver strategy decisions. This latter application was the first use of Monte Carlo trajectory analysis in an aerobraking 

mission. 

Author 

Aerobraking; Capture Effect; Flight Operations; Navigation; Planning; Simulation; Trajectories 

20030005452 NASA Johnson Space Center, Houston, TX USA 
Aerocapture Guidance Algorithm Comparison Campaign 

Rousseau, Stephane; Perot, Etienne; Graves, Claude; Masciarelli, James P.; Queen, Eric; [2002]; In English, 5-8 Aug. 2002, 
Monterey, CA, USA; Original contains color illustrations 

Report No.(s): AIAA Paper 2002-4822; Copyright; Avail: CASI; A02, Hardcopy; Distribution as joint owner in the copyright 
The aerocapture is a promising technique for the future human interplanetary missions. The Mars Sample Return was 
initially based on an insertion by aerocapture. A CNES orbiter Mars Premier was developed to demonstrate this concept. 
Mainly due to budget constraints, the aerocapture was cancelled for the French orbiter. A lot of studies were achieved during 
the three last years to develop and test different guidance algorithms (APC, EC, TPC, NPC). This work was shared between 
CNES and NASA, with a fruitful joint working group. To finish this study an evaluation campaign has been performed to test 
the different algorithms. The objective was to assess the robustness, accuracy, capability to limit the load, and the complexity 
of each algorithm. A simulation campaign has been specified and performed by CNES, with a similar activity on the NASA 
side to confirm the CNES results. This evaluation has demonstrated that the numerical guidance principal is not competitive 
compared to the analytical concepts. All the other algorithms are well adapted to guaranty the success of the aerocapture. The 
TPC appears to be the more robust, the APC the more accurate, and the EC appears to be a good compromise. 
Author 
Aerocapture; Spacecraft Guidance; Algorithms 

20030005447 NASA Langley Research Center, Hampton, VA USA 

Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap 

Horvath, Thomas J.; OConnell, Tod F; Cheatwood, F. McNeil; Prabhu, Ramadas K.; Alter, Stephen J.; [2002]; In English, 5-8 
Aug. 2002, Monterey, CA, USA 

Report No.(s): AIAA Paper 2002-4408; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright 
Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001 
Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective 
was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. 
The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were 
obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody 
diameter) ranging from 2,637 to 92,350, angles of attack ranging from tip to 23 degrees at and 2 degree sideslip, and 
normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy 
gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. 
The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of 
planetary entry vehicle are not possible (e.g. offset e.g.), a single flap can provide similar aerodynamic performance. An 
assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air 
at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide 
sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured 
grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were 
found to be in very good agreement with experimental measurement. 
Author 
Aerodynamic Characteristics; Aerothermodynamics; Hypersonic Flow; Inviscid Flow 

20030002240 NASA Langley Research Center, Hampton, VA USA 
Computational Analysis of Towed Ballute Interactions 

Gnoffo, Peter A.; Anderson, Brian P.; [2002]; In English, 24-26 Jun. 2002, Saint Louis, MO, USA; Original contains color 

illustrations 

Report No.(s): AIAA Paper 2002-2997; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright 

18 



A ballute (balloon-parachute) is an inflatable, aerodynamic drag device for application to planetary entry vehicles. 
Ballutes may be directly attached to a vehicle, increasing its cross-sectional area upon inflation, or towed behind the vehicle 
as a semi-independent device that can be quickly cut free when the requisite change in velocity is achieved. The 
aero thermodynamics of spherical and toroidal towed ballutes are considered in the present study. A limiting case of zero 
towline length (clamped system) is also considered. A toroidal system can be designed (ignoring influence of the tethers) such 
that all flow processed by the bow shock of the towing spacecraft passes through the hole in the toroid. For a spherical ballute, 
towline length is a critical parameter that affects aeroheating on the ballute being towed through the spacecraft wake. In both 
cases, complex and often unsteady interactions ensue in which the spacecraft and its wake resemble an aero spike situated in 
front of the ballute. The strength of the interactions depends upon system geometry and Reynolds number. We show how 
interactions may envelope the base of the towing spacecraft or impinge on the ballute surface with adverse consequences to 
its thermal protection system. Geometric constraints to minimize or eliminate such adverse interactions are discussed. The 
towed, toroidal system and the clamped, spherical system show greatest potential for a baseline design approach. 
Author 
Atmospheric Entry; Ballutes; Spacecraft Control; Flight Control; Towed Bodies; Computerized Simulation 

20030002226 NASA Langley Research Center, Hampton, VA USA 

Application of Accelerometer Data to Mars Odyssey Aerobraking and Atmospheric Modeling 

Tolson, R. H.; Keating, G. M.; George, B. E.; Escalera, P. E.; Werner, M. R.; Dwyer, A. M.; Hanna, J. L.; [2002]; In English, 
5-8 Aug. 2002, Monterey, CA, USA; Original contains color illustrations 

Report No.(s): AIAA Paper 2002-4533; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright 
Aerobraking was an enabling technology for the Mars Odyssey mission even though it involved risk due primarily to the 
variability of the Mars upper atmosphere. Consequently, numerous analyses based on various data types were performed 
during operations to reduce these risk and among these data were measurements from spacecraft accelerometers. This paper 
reports on the use of accelerometer data for determining atmospheric density during Odyssey aerobraking operations. 
Acceleration was measured along three orthogonal axes, although only data from the component along the axis nominally into 
the flow was used during operations. For a one second count time, the RMS noise level varied from 0.07 to 0.5 mm/s2 
permitting density recovery to between 0.15 and 1.1 kg per cu km or about 2% of the mean density at periapsis during 
aerobraking. Accelerometer data were analyzed in near real time to provide estimates of density at periapsis, maximum 
density, density scale height, latitudinal gradient, longitudinal wave variations and location of the polar vortex. Summaries are 
given of the aerobraking phase of the mission, the accelerometer data analysis methods and operational procedures, some 
applications to determining thermospheric properties, and some remaining issues on interpretation of the data. Pre-flight 
estimates of natural variability based on Mars Global Surveyor accelerometer measurements proved reliable in the 
mid-latitudes, but overestimated the variability inside the polar vortex. 
Author 

Accelerometers; 2001 Mars Odyssey; Aerobraking; Atmospheric Models; Mars Atmosphere; Atmospheric Density; Radio 
Tracking; Numerical Analysis; Trajectory Analysis 

20030000829 NASA Langley Research Center, Hampton, VA USA 
Mars Smart Lander Parachute Simulation Model 

Queen, Eric M.; Raiszadeh, Ben; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA 

Report No.(s): AIAA Paper 2002-4617; Copyright; Avail: CASI; A02, Hardcopy; Distribution under U.S. Government purpose 

rights 

A multi-body flight simulation for the Mars Smart Lander has been developed that includes six degree-of-freedom 
rigid-body models for both the supersonically-deployed and subsonically-deployed parachutes. This simulation is designed to 
be incorporated into a larger simulation of the entire entry, descent and landing (EDL) sequence. The complete end-to-end 
simulation will provide attitude history predictions of all bodies throughout the flight as well as loads on each of the connecting 
lines. Other issues such as recontact with jettisoned elements (heat shield, back shield, parachute mortar covers, etc.), design 
of parachute and attachment points, and desirable line properties can also be addressed readily using this simulation. 
Author 
Flight Simulation; Parachutes; Mars Landing; Trajectory Analysis; Evaluation; Loads (Forces) 

20030000735 CFD Research Corp., Huntsville, AL USA 

CFD Prediction of the BEAGLE 2 Mars Probe Aerodynamic Database 

Liever, Peter A.; Habchi, Sami D.; Burnell, Simon I.; Lingard, Steve J.; Twelfth Thermal and Fluids Analysis Workshop; July 
2002; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy 

19 



CFD (Computational Fluid Dynamics) has matured to provide reliable planetary entry vehicle aerodynamic predictions. 
CFD provides substantial time and cost savings. CFD-FASTRAN was applied over the entire trajectory (Entry to Chute 
Deployment). It provided valuable insight into vehicle flow characteristics (Examples: Wake and Base Flow Structure, 
Transonic Wake Unsteadiness). A blended aerodynamic database was generated by combining CFD data, scaled existing data, 
and wind tunnel test data. CFD based pitch damping analysis provides insight into dynamic stability characteristics not easily 
obtained from wind tunnel tests. 
Derived from text 
Computational Fluid Dynamics; Wind Tunnel Tests; Atmospheric Entry; Predictions; Data Bases 

20020081342 NASA Glenn Research Center, Cleveland, OH USA 
Radioisotope Electric Propulsion for Fast Outer Planetary Orbiters 

Oleson, Steven; Benson, Scott; Gefert, Leon; Patterson, Michael; Schreiber, Jeffrey; September 2002; In English, 7-10 Jul. 

2002, Indianapolis, IN, USA 

Contract(s)/Grant(s): RTOP 344-96-8D 

Report No.(s): NASA/TM-2002-211893; NAS 1.15:211893; E-13575; AIAA Paper 2002-3967; No Copyright; Avail: CASI; 

A03, Hardcopy 

Recent interest in outer planetary targets by the Office of Space Science has spurred the search for technology options to 
enable relatively quick missions to outer planetary targets. Several options are being explored including solar electric propelled 
stages combined with aerocapture at the target and nuclear electric propulsion. Another option uses radioisotope powered 
electric thrusters to reach the outer planets. Past work looked at using this technology to provide faster fiybys. A better use 
for this technology is for outer planet orbiters. Combined with medium class launch vehicles and a new direct trajectory these 
small, sub-kilowatt ion thrusters and Stirling radioisotope generators were found to allow missions as fast as 5 to 12 years for 
objects from Saturn to Pluto, respectively. Key to the development is light spacecraft and science payload technologies. 
Author 
Nuclear Electric Propulsion; Radioactive Isotopes; Gas Giant Planets; Grand Tours; Aerocapture 

20020077966 NASA Ames Research Center, Moffett Field, CA USA 
Aerocapture Technology Development Needs for Outer Planet Exploration 

Wercinski, Paul; Munk, Michelle; Powell, Richard; Hall, Jeff; Graves, Claude; Partridge, Harry, Technical Monitor; January 

2002; In English 

Contract(s)/Grant(s): RTOP 713-81-80 

Report No.(s): NASA/TM-2002-211386; NAS 1.15:211386; A-0107378; No Copyright; Avail: CASI; A03, Hardcopy 

The purpose of this white paper is to identify aerocapture technology and system level development needs to enable 
NASA future mission planning to support Outer Planet Exploration. Aerocapture is a flight maneuver that takes place at very 
high speeds within a planet's atmosphere that provides a change in velocity using aerodynamic forces (in contrast to 
propulsive thrust) for orbit insertion. Aerocapture is very much a system level technology where individual disciplines such 
as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS), 
guidance, navigation and control (GN&C) instrumentation need to be integrated and optimized to meet mission specific 
requirements. This paper identifies on-going activities, their relevance and potential benefit to outer planet aerocapture that 
include New Millennium ST7 Aerocapture concept definition study, Mars Exploration Program aeroassist project level 
support, and FY01 Aeroassist In-Space Guideline tasks. The challenges of performing aerocapture for outer planet missions 
such as Titan Explorer or Neptune Orbiter require investments to advance the technology readiness of the aerocapture 
technology disciplines for the unique application of outer planet aerocapture. This white paper will identify critical technology 
gaps (with emphasis on aeroshell concepts) and strategies for advancement. 
Author 
Aerocapture; Aerothermodynamics; Spacecraft Design; Orbit Insertion; Spacecraft Maneuvers; Outer Planets Explorers 

20020039836 NASA Langley Research Center, Hampton, VA USA 

Aerothermal Instrumentation Loads To Implement Aeroassist Technology in Future Robotic and Human Missions to 

MARS and Other Locations Within the Solar System 

Parmar, Devendra S.; Shams, Qamar A.; April 2002; In English 

Contract(s)/Grant(s): RTOP 713-81-70 

Report No.(s): NASA/TM-2002-211459; NAS 1.15:211459; L-18123; No Copyright; Avail: CASI; A03, Hardcopy 

20 



The strategy of NASA to explore space objects in the vicinity of Earth and other planets of the solar system includes 
robotic and human missions. This strategy requires a road map for technology development that will support the robotic 
exploration and provide safety for the humans traveling to other celestial bodies. Aeroassist is one of the key elements of 
technology planning for the success of future robot and human exploration missions to other celestial bodies. Measurement 
of aerothermodynamic parameters such as temperature, pressure, and acceleration is of prime importance for aeroassist 
technology implementation and for the safety and affordability of the mission. Instrumentation and methods to measure such 
parameters have been reviewed in this report in view of past practices, current commercial availability of instrumentation 
technology, and the prospects of improvement and upgrade according to the requirements. Analysis of the usability of each 
identified instruments in terms of cost for efficient weight-volume ratio, power requirement, accuracy, sample rates, and other 
appropriate metrics such as harsh environment survivability has been reported. 
Author 

Aeroassist; Aerothermodynamics; Robotics; Technology Utilization; Aerodynamic Loads; Solar System; Manned Mars 
Missions; Temperature Measuring Instruments 

20020023456 Instituto Nacional de Pesquisas Espacias, Sao Jose dos Campos, Brazil 
Study of Orbital Transfers with Aeroassisted Maneuvers 

Schulz, Walkiria; 2001; In Portuguese 

Report No.(s): INPE-8476-TDI/776; No Copyright; Avail: CASI; A09, Hardcopy 

This work presents an analysis of space missions through the development of a software package for the calculation of 
aerodynamic maneuvers and of the required thrust maneuvers for their implementation. Besides the numerical development, 
an analytical study contemplates the accomplishment of the aeroassisted phase of this maneuver type. This analysis includes 
a study of the thermal limits associated with a vehicle passage through the atmosphere as well as an analysis of the associated 
errors. Several simulations of aerodynamic maneuvers are carried out and compared with orbital changes accomplished 
outside of the atmosphere. Among the conclusions, it is shown that the problem is extremely sensitive to the initial conditions 
and each mission deserves a careful individual analysis. Finally, the results obtained from the analytical formulation are shown 
to be in agreement with the numerical results for the upper layers of the terrestrial atmosphere. 
Author 
Transfer Orbits; Space Missions; Applications Programs (Computers); Aerodynamic Characteristics; Aeroassist 

20020002105 NASA Johnson Space Center, Houston, TX USA 
Beagle 2: The Next Exobiology Mission to Mars 

Gibson, Everett K., Jr.; Pillinger, Colin T; Wright, Ian P.; Morse, Andy; Stewart, Jenny; Morgan, G.; Praine, Ian; Leigh, 
Dennis; Sims, Mark R.; General Meeting of the NASA Astrobiology Insititute; April 2001, 160-162; In English; No 
Copyright; Avail: Other Sources; Abstract Only 

Beagle 2 is a 60 kg probe (with a 30 kg lander) developed in the UK for inclusion on the European Space Agency's 2003 
Mars Express. Beagle 2 will deliver to the Martian surface a payload which consists of a high percentage of science 
instruments to landed spacecraft mass. Beagle 2 will be launched in June, 2003 with Mars Express on a Soyuz-Fregat rocket 
from the Baikonur Cosmodrome in Kazakhstan. Beagle 2 will land on Mars on December 26, 2003 in the Isidis Planitia basin 
(approximately 10 degrees N and 275 degrees W), a large sedimentary basin that overlies the boundary between ancient 
highlands and northern plains. Isidis Planitia, the third largest basin on Mars, which is possibly filled with sediment deposited 
at the bottom of long-standing lakes or seas, offers an ideal environment for preserving traces of life. Beagle 2 was developed 
to search for organic material and other volatiles on and below the surface of Mars in addition to the study of the inorganic 
chemistry and mineralogy. Beagle 2 will utilize a mechanical mole and grinder to obtain samples from below the surface, 
under rocks and inside rocks. A pair of stereo cameras will image the landing site along with a microscope for examination 
of surface and rock samples. Analyses will include both rock and soil samples at various wavelengths, X-ray spectrometer and 
Mossbauer spectrometer as well as a search for organics and other light element species (e.g. carbonates and water) and 
measurement of their isotopic compositions. Beagle 2 has as its focus the goal of establishing whether evidence for life existed 
in the past on Mars at the Isidis Planitia site or at least establishing if the conditions were ever suitable. Carbonates and organic 
components were first recognized as existing on Mars when they were found in the Martian meteorite Nakhla. Romanek et 
al showed the carbonates in ALH84001 were formed at low temperatures. McKay et al noted possible evidence of early life 
on Mars within the ALH84001 meteorite. Thomas-Keprta et al showed the magnetite biomarkers in ALH84001's carbonates 
are indistinguishable from those formed by magnetotactic bacteria found on Earth. Gibson et al showed there was significant 
evidence for liquid water and biogenic products present on Mars across a 3.9 billion year period. A mechanical arm (PAW) 
operates from the lander and is used for science operations along with sample acquisition. Instruments attached to the PAW 

21 



include: stereo cameras, Mossbauer instrument, X-ray fluorescence instrument, microscope, environmental sensors, rock 
corer/grinder, a spoon, mirror, brushes, a mole attachment for acquisition of subsurface to depths of 1 to 2 meters and an 
illumination device. Each camera has 14 filters which have been optimized for mineralogy composition, dust and water vapor 
detection. The microscope's camera is designed for viewing the size and shape of dust particles, rock surfaces, microfossils, 
and characteristics of the samples prior to introduction into the gas analysis package (GAP). The camera has a resolution of 
4 microns/pixel, features four color capability (red, green, blue and UV (ultraviolet) fluoresence), a depth of focus of 40 
micrometers and translation stage of +3 millimeters. The heart of the Beagle 2's life detection package is the gas analysis 
package (GAP), which consists of a mass spectrometer with collectors at fixed masses for precise isotopic ratio measurements 
and voltage scanning for spectral analysis. Primary aim of the GAP is to search for the presence of bulk constituents, individual 
species, and isotopic fractionations for both extinct and extent life along with studying the low-temperature geochemistry of 
the hydrogen, carbon, nitrogen and oxygen components on Mars from both the surface and atmosphere. GAP is a magnetic 
sector mass spectrometer with the range of 1 to 140 amu which can be operated in both the static and dynamic modes. A triple 
Faraday collector array will be used for C, N and O ratios along with a double Faraday array for H/D. Pulse counting electron 
multiplier will be utilized for noble gases and selected organics. Anticipated detection limits are at the picomole level for 
operation in the static mode of operation and high precision isotopic measurements will be made in the dynamic mode. Sample 
processing and preparation system consists of reaction vessels along with references. Sample ovens capable of being heated 
are attached to the manifold for sample combustion. Surface, subsurface materials and interior rock specimens will be 
combusted in pure oxygen gas at various temperature intervals to release organic matter and volatiles. Combustion process 
will permit detection of all forms and all atoms of carbon present in the samples. A chemical processing system is capable of 
a variety of conversion reactions. Gases are manipulated either by cryogenic or chemical reactions and passed through the gas 
handling portion of the vacuum system. There are two modes of operation: quantitative analysis and precise isotopic 
measurements. Additional information is contained in the original extended abstract. 
Author 
Exobiology; Mars Missions; Spacecraft Instruments; European Space Programs; Space Probes 



20010122748 Tennessee Univ., Knoxville, TN USA 

Remote Sensing of Evaporite Minerals in Badwater Basin, Death Valley, at Varying Spatial Scales and in Different 

Spectral Regions 

Moersch, J. E.; Farmer, J.; Baldridge, A.; Field Trip and Workshop on the Martian Highlands and Mojave Desert Analogs; 
2001, 45-46; In English; No Copyright; Avail: CASI; A01, Hardcopy 

A key concept behind the overall architecture of NASA's Mars Surveyor Program is that remote sensing observations 
made from orbit will be used to guide the selection of landing sites for subsequent missions to the surface. An important 
component of the orbital phase of this strategy is mineralogical mapping of the surface with infrared spectrometers and 
imaging systems. Currently, the Mars Global Surveyor Thermal Emission Spectrometer (TES) is spectrally mapping Mars in 
the 6-50 micrometer region at a spatial resolution of 3 km. Starting later this year, the Thermal Emission Imaging System 
(THEMIS) aboard the Mars 2001 Odyssey orbiter will image the entire surface of the planet in eight broad bands in the 
6.5-14.5 micrometer region at a spatial resolution of 100 m. In 2003, ESA plans to launch the OMEGA instrument on Mars 
Express, which will map the planet in the visible and near infrared regions from an elliptical orbit at spatial resolutions of up 
to 100 m. Currently, NASA is selecting a visible and near-infrared mapping spectrometer for an orbiter that will launch in 
2005. This instrument will likely map at a constant spatial resolution of at least 50 m. From an astrobiological perspective, 
the utility of these spectral datasets will be in locating potential paleohabitats for martian life, via the detection of minerals 
that form in the presence of liquid water. Deposits of evaporite minerals in putative martian paleolake basins are a particularly 
attractive target to look for because of the areal extent of these features, the strong spectral features of these minerals, and the 
characteristic sequences in which they appear along the margin of a basin. Despite considerable geomorphic evidence 
indicating the presence of ancient lake basins on Mars, to date no evaporite deposits have been reported from the TES 
experiment. But is this to be expected, given the limited spatial resolution of TES data? Might we still hope to find such 
deposits in upcoming experiments? One way to address this question is to use existing datasets from terrestrial analog sites 
to attempt to determine spatial and spectral thresholds of detectability for these minerals in a natural setting 
Author 

Remote Sensing; Death Valley (CA); Spatial Resolution; Spectral Bands; Mineral Deposits; Structural Basins; Sedimentary 
Rocks 

22 



20010099686 Georgia Inst, of Tech., Atlanta, GA USA 

Uncertainty Optimization Applied to the Monte Carlo Analysis of Planetary Entry Trajectories 

Olds, John; Way, David; Jul. 31, 2001; In English 

Contract(s)/Grant(s): NGT1-52163; No Copyright; Avail: CASI; All, Hardcopy 

Recently, strong evidence of liquid water under the surface of Mars and a meteorite that might contain ancient microbes 
have renewed interest in Mars exploration. With this renewed interest, NASA plans to send spacecraft to Mars approx. every 
26 months. These future spacecraft will return higher-resolution images, make precision landings, engage in longer-ranging 
surface maneuvers, and even return Martian soil and rock samples to Earth. Future robotic missions and any human missions 
to Mars will require precise entries to ensure safe landings near science objective and pre-employed assets. Potential sources 
of water and other interesting geographic features are often located near hazards, such as within craters or along canyon walls. 
In order for more accurate landings to be made, spacecraft entering the Martian atmosphere need to use lift to actively control 
the entry. This active guidance results in much smaller landing footprints. Planning for these missions will depend heavily on 
Monte Carlo analysis. Monte Carlo trajectory simulations have been used with a high degree of success in recent planetary 
exploration missions. These analyses ascertain the impact of off-nominal conditions during a flight and account for uncertainty. 
Uncertainties generally stem from limitations in manufacturing tolerances, measurement capabilities, analysis accuracies, and 
environmental unknowns. Thousands of off-nominal trajectories are simulated by randomly dispersing uncertainty variables 
and collecting statistics on forecast variables. The dependability of Monte Carlo forecasts, however, is limited by the accuracy 
and completeness of the assumed uncertainties. This is because Monte Carlo analysis is a forward driven problem; beginning 
with the input uncertainties and proceeding to the forecasts outputs. It lacks a mechanism to affect or alter the uncertainties 
based on the forecast results. If the results are unacceptable, the current practice is to use an iterative, trial-and-error approach 
to reconcile discrepancies. Therefore, an improvement to the Monte Carlo analysis is needed that will allow the problem to 
be worked in reverse. In this way, the largest allowable dispersions that achieve the required mission objectives can be 
determined quantitatively. 
Derived from text 

Atmospheric Entry; Trajectory Optimization; Monte Carlo Method; Trajectory Planning; Trajectory Analysis; Trajectory 
Control; Mathematical Models 

20010041296 NASA Ames Research Center, Moffett Field, CA USA 
Exploration of Titan Using Vertical Lift Aerial Vehicles 

Young, L. A.; Forum on Innovative Approaches to Outer Planetary Exploration 2001-2020; 2001, 94; In English; No 
Copyright; Abstract Only; Available from CASI only as part of the entire parent document 

Autonomous vertical lift aerial vehicles (such as rotorcraft or powered-lift vehicles) hold considerable potential for 
supporting planetary science and exploration missions. Vertical lift aerial vehicles would have the following advantages/ 
attributes for planetary exploration: (1) low-speed and low-altitude detailed aerial surveys; (2) remote-site sample return to 
lander platforms; (3) precision placement of scientific probes; (4) soft landing capability for vehicle reuse (multiple flights) 
and remote-site monitoring; (5) greater range, speed, and access to hazardous terrain than a surface rover; and (6) greater 
resolution of surface details than an orbiter or balloons. Exploration of Titan presents an excellent opportunity for the 
development and usage of such vehicles. Additional information is contained in the original extended abstract. 
Derived from text 
Vertical Takeoff Aircraft; Space Exploration; Powered Lift Aircraft; Rotary Wing Aircraft 

20010041208 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
Micro Navigator 

Blaes, B. R.; Kia, T; Chau, S. N; Forum on Innovative Approaches to Outer Planetary Exploration 2001-2020; 2001, 9; In 
English; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document 

Miniature high-performance low-mass space avionics systems are desired for planned future outer planetary exploration 
missions (i.e. Europa Orbiter/Lander, Pluto-Kuiper Express). The spacecraft fuel and mass requirements enabling orbit 
insertion is the driving requirement. The Micro Navigator is an integrated autonomous Guidance, Navigation & Control 
(GN&C)micro-system that would provide the critical avionics function for navigation, pointing, and precision landing. The 
Micro Navigator hardware and software allow fusion of data from multiple sensors to provide a single integrated vehicle state 
vector necessary for six degrees of freedom GN&C. The benefits of this MicroNavigator include: 1) The Micro Navigator 
employs MEMS devices that promise orders of magnitude reductions in mass power and volume of inertial sensors 
(accelerometers and gyroscopes), celestial sensing devices (startracker, sun sensor), and computing element; 2) The highly 
integrated nature of the unit will reduce the cost of flight missions, a) The advanced miniaturization technologies employed 

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by the Micro Navigator lend themselves to mass production, and therefore will reduce production cost of spacecraft, b) The 
integral approach simplifies interface issues associated with discrete components and reduces cost associated with integration 
and test of multiple components; and 3) The integration of sensors and processing elements into a single unit will allow the 
Micro Navigator to encapsulate attitude information and determination functions into a single object. This is particularly 
beneficial for object-oriented software architectures that are used in advanced spacecraft. Additional information is contained 
in the original extended abstract. 
Derived from text 
Autonomous Navigation; Miniaturization; Multisensor Fusion; Space Navigation; Spacecraft Guidance 

20010038566 NASA Ames Research Center, Moffett Field, CA USA 
International Agreement on Planetary Protection 

Mars Sample Handling Protocol Workshop Series; October 2000, 93; In English; No Copyright; Avail: CASI; A01, Hardcopy 
The maintenance of a NASA policy, is consistent with international agreements. The planetary protection policy 
management in OSS, with Field Center support. The advice from internal and external advisory groups (NRC, NAC/Planetary 
Protection Task Force). The technology research and standards development in bioload characterization. The technology 
research and development in bioload reduction/sterilization. This presentation focuses on: forward contamination - research 
on the potential for Earth life to exist on other bodies, improved strategies for planetary navigation and collision avoidance, 
and improved procedures for sterile spacecraft assembly, cleaning and/or sterilization; and backward contamination - 
development of sample transfer and container sealing technologies for Earth return, improvement in sample return landing 
target assessment and navigation strategy, planning for sample hazard determination requirements and procedures, safety 
certification, (liaison to NEO Program Office for compositional data on small bodies), facility planning for sample recovery 
system, quarantine, and long-term curation of 4 returned samples. 
Derived from text 
International Cooperation; International Law; Policies; Planetary Environments; Environment Protection 

20010024974 Colorado Univ., Boulder, CO USA 
Low Velocity Impact Experiments in Microgravity 

Colwell, J. E.; Sture, S.; Proceedings of the Fifth Microgravity Fluid Physics and Transport Phenomena Conference; 
December 2000, 1335-1346; In English; No Copyright; Avail: CASI; A03, Hardcopy 

Protoplanetary disks, planetary rings, the Kuiper belt, and the asteroid belt are collisionally evolved systems. Although 
objects in each system may be bombarded by impactors at high interplanetary velocities, they are also subject to repeated 
collisions at low velocities (v\hl0 m/s). In some regions of Saturn's rings, for example, the typical collision velocity inferred 
from observations by the Voyager spacecraft and dynamical modeling is a fraction of a centimeter per second. These 
interparticle collisions control the rate of energy dissipation in planetary rings and the rate of accretion in the early stages of 
planetesimal formation. In the asteroid belt collisions typically occur at several km/s; however secondary craters are formed 
at much lower impact speeds. In the Kuiper belt, where orbital speeds and eccentricities are much lower, collisions between 
Kuiper belt objects (KBOs) can occur at speeds below 100 m/s. In the early solar system, KBOs accreted in the same way 
planetesimals accreted in the inner solar system, however some regions of the Kuiper belt may now undergo erosional 
collisions. Dust may be present on the surface of all of these objects in the form of a fine regolith created from micrometeoroid 
bombardment (rings, asteroids, KBOs), high speed interparticle collisions (asteroids, KBOs) or as a product of accretion from 
protoplanetary dust. Dust released in these collisions is often the only observable trace of the source objects and may be used 
to infer the physical properties of those larger bodies. We are conducting a broad program of microgravity impact experiments 
into dust to study the dissipation of energy in low energy collisions and the production of dust ejecta in these impacts. The 
Collisions Into Dust Experiment (COLLIDE) flew on STS-90 in April 1998. The principal results of that experiment were 
measurements of the coefficient of restitution for impacts into powders at impact speeds below 1 m/s. Almost no ejecta was 
produced in impacts at 15 cm/s into JSC-1 powder, and the coefficient of restitution was about 0.03. COLLIDE-2 is undergoing 
final preparations for a flight in 2001. The experiment will conduct six impact experiments at impact speeds between 1 and 
100 cm/s. The target material will have a low relative density to mimic the regolith on low surface gravity objects in space, 
such as planetesimals, planetary ring particles, and asteroids. A new experimental program, the Physics of Regolith Impacts 
in Microgravity Experiment (PRIME) will use NASA's KC-135 aircraft to explore a much broader range of parameter space 
than is possible with COLLIDE, at slightly higher impact velocities. PRIME will be capable of conducting up to 16 impact 
experiments each flight day on the KC-135. Impact velocities between 50 cm/s and 5 m/s will be studied into a variety of target 
materials and size distributions. The experiment will consist of an evacuated canister with 6 to 8 impact chambers on each 
of two rotating turntables. Each impact chamber will include a target sample and a launcher with a unique set of parameters. 

24 



Two viewports will allow high speed video photography of impacts from two orthogonal views with the use of a mirror 
mounted inside the canister. Data from COLLIDE and ground-based experimental studies suggest that particle size distribution 
is an important parameter in controlling the response of granular media to low velocity impacts. Individual grain shapes may 
also play an important role in the conversion of impactor kinetic energy to target grain kinetic energy. We will also make use 
of numerical simulations of the impact process to understand the relevant parameters for experimental study. High speed video 
of the impact and ejecta patterns will be analyzed to determine the ejecta mass and velocity distributions. This in turn will have 
direct application for understanding the behavior of dust on the surfaces of planetary objects including asteroids and small 
moons when disturbed by low velocity impacts and perturbations. These include naturally occurring impacts as well as 
disturbances to the surface from human and spacecraft activity. The velocity distribution of the ejecta determines the amount 
of material launched to various altitudes above the surface and escaping the parent body. This information is important for 
spacecraft instruments landing on airless bodies with low surface gravity and powdery regoliths. 
Author (revised) 
Collisions; Dust; Gravitational Effects; Low Speed; Microgravity; Impact; Energy Dissipation 

20010023145 Centre National d'Etudes Spatiales, France 
The Stakes of the Aerocapture for Missions to Mars 

Cledassou, R.; Lam-Trong, Th.; Charbonnier, J. M.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 1, 
186; In English; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document 

The Hohmann transfer trajectory is an economical way to go from Earth to Mars but a spacecraft has to reduce its speed 
very significantly upon arrival in order to be inserted into a Mars orbit. The aerocapture is a way to do that, by using the 
Martian atmosphere to produce sufficient aerodynamic drag force on a heatshield and achieve the required deceleration. This 
presentation will address the major stake of the aerocapture which is twofold: a) We will list the different technologies and 
areas of knowledge related to the aerocapture, identify the risks associated with each of them and finally demonstrate that 
aerocapture is not as risky as it is said to be; b) Aerocapture saves a huge amount of propellant and so allows to improve 
dramatically the dollar/kg ratio for any payload at Mars by using this mass savings for payloads and by decreasing the launch 
cost. This benefit is particularly evident for a return mission because of the amplification factor of the propellant mass for the 
escape of Mars ('snow ball' effect). We will have a quantitative analysis of some typical cases of spacecraft vs. launcher 
performance . We will conclude that aerocapture is interesting for the present robotic missions and certainly a good investment 
for the future manned missions to Mars. 
Derived from text 
Mars Missions; Aerocapture; Earth-Mars Trajectories; Transfer Orbits 

20010023141 Science Applications International Corp., Littleton, CO USA 
Precision Terminal Guidance for a Mars Lander 

Klarquist, William N.; Wahl, Beth E.; Lowrie, James W.; Concepts and Approaches for Mars Exploration; July 2000, Issue 
Part 1, 178-179; In English; No Copyright; Avail: CASI; A01, Hardcopy 

To date Mars landers have relied solely on Earth-based navigation measurements to achieve a desired landing site. 
They've had no active guidance and control system to monitor and control the entry and descent trajectory or guide the final 
landing. This results in very large landing site uncertainties (\gl80 km x 20 km) and precludes targeting specific, small scale 
regions such as canyons and flood channels. Moreover, localized hazards cannot be sensed or avoided, resulting in higher 
mission risk. SAIC's Center for Intelligent Systems, (SAIC-CIS) based on current and past research, believes that reliably 
accurate landings at pre-selected sites are achievable and that the mission risk associated with local hazards can be greatly 
reduced. Our concept involves applying an integrated system level solution that leverages the tremendous amount of 
information available on the Martian environment and applies modern technologies in the areas of visual based navigation, 
maneuverable parachutes, and advanced sensors. 
Derived from text 
Active Control; Terminal Guidance; Spacecraft Guidance; Mars Landing 

20010023136 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
Mars Sample Return without Landing on the Surface 

Jurewicz, A. J. G.; Jones, Steven M.; Yen, A. S.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 1, 
168-169; In English; No Copyright; Avail: CASI; A01, Hardcopy 

Many in the science community want a Mars sample return in the near future, with the expectation that it will provide 

25 



in-depth information, significantly beyond what we know from remote sensing, limited in-situ measurements, and work with 
Martian meteorites. Certainly, return of samples from the Moon resulted in major advances in our understanding of both the 
geologic history of our planetary satellite, and its relationship to Earth. Similar scientific insights would be expected from 
analyses of samples returned from Mars. Unfortunately, Mars-lander sample-return missions have been delayed, for the reason 
that NASA needs more time to review the complexities and risks associated with that type of mission. A traditional sample 
return entails a complex transfer-chain, including landing, collection, launch, rendezvous, and the return to Earth, as well as 
an evaluation of potential biological hazards involved with bringing pristine Martian organics to Earth. There are, however, 
means of returning scientifically-rich samples from Mars without landing on the surface. This paper discusses an approach for 
returning intact samples of surface dust, based on known instrument technology, without using an actual Martian lander. 
Derived from text 
Mars Sample Return Missions; Particulate Sampling; Aerocapture 

20010020516 Arizona Univ., Tucson, AZ USA 
The Martian Oasis Detector 

Smith, P. H.; tomasko, M. G.; McEwen, A.; Rice, J.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 
2, 286-287; In English; No Copyright; Avail: CASI; A01, Hardcopy 

The next phase of unmanned Mars missions paves the way for astronauts to land on the surface of Mars. There are lessons 
to be learned from the unmanned precursor missions to the Moon and the Apollo lunar surface expeditions. These unmanned 
missions (Ranger, Lunar Orbiter, and Surveyor) provided the following valuable information, useful from both a scientific and 
engineering perspective, which was required to prepare the way for the manned exploration of the lunar surface: (1) high 
resolution imagery instrumental to Apollo landing site selection also tremendously advanced the state of Nearside and Farside 
regional geology; (2) demonstrated precision landing (less than two kilometers from target) and soft landing capability; (3) 
established that the surface had sufficient bearing strength to support a spacecraft; and (4) examination of the chemical 
composition and mechanical properties of the surface. The search for extinct or extant life on Mars will follow the water. 
However, geomorphic studies have shown that Mars has had liquid water on its surface throughout its geologic history. A 
cornucopia of potential landing sites with water histories (lakes, floodplains, oceans, deltas, hydrothermal regions) presently 
exist. How will we narrow down site selection and increase the likelihood of finding the signs of life? One way to do this is 
to identify 'Martian oases.' It is known that the Martian surface is often highly fractured and some areas have karst structures 
that support underground caves. Much of the water that formed the channels and valley networks is thought to be frozen 
underground. All that is needed to create the potential for liquid water is a near surface source of heat; recent lava flows and 
Martian meteorites attest to the potential for volcanic activity. If we can locate even one spot where fracturing, ice, and 
underground heat are co-located then we have the potential for an oasis. Such a discovery could truly excite the imaginations 
of both the public and Congress providing an attainable goal for both robotic and manned missions. The instrument required 
to detect an active oasis is a high spatial resolution (few tens of meters) Short Wavelength Infrared (SWIR) spectrometer 
coupled with a high resolution camera (five m/pixel). This combination creates too large a data volume to possibly return data 
for the entire Martian Surface; therefore it has been designed as one of the first in a new generation of 'smart' detectors, called 
the Mars Oasis Detector (MOD). 
Author 
Mars (Planet); Water; Mars Missions; Mars Surface; Spacecraft Instruments; Oases; Mars Exploration 

20010020495 California Univ., Los Angeles, CA USA 
After the Mars Polar Lander: Where to Next? 

Paige, D. A.; Boynton, W. V.; Crisp, D.; DeJong, E.; Hansen, C. J.; Harri, A. M.; Keller, H. U.; Leshin, L. A.; May, R. D.; 
Smith, P. H., et al.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 2, 245-246; In English; No 
Copyright; Avail: CASI; A01, Hardcopy 

The recent loss of the Mars Polar Lander (MPL) mission represents a serious setback to Mars science and exploration. 
Targeted to land on the Martian south polar layered deposits at 76 degrees south latitude and 195 degrees west longitude, it 
would have been the first mission to study the geology, atmospheric environment, and volatiles at a high-latitude landing site. 
Since the conception of the MPL mission, a Mars exploration strategy has emerged which focuses on Climate, Resources and 
Life, with the behavior and history of water as the unifying theme. A successful MPL mission would have made significant 
contributions towards these goals, particularly in understanding the distribution and behavior of near-surface water, and the 
nature and climate history of the south polar layered deposits. Unfortunately, due to concerns regarding the design of the MPL 
spacecraft, the rarity of direct trajectories that enable high-latitude landings, and funding, an exact reflight of MPL is not 
feasible within the present planning horizon. However, there remains significant interest in recapturing the scientific goals of 

26 



the MPL mission. The following is a discussion of scientific and strategic issues relevant to planning the next polar lander 

mission, and beyond. 

Author 

Mars Exploration; Mission Planning; Mars Missions; Polar Regions; Mars (Planet) 

20010020467 Science Applications International Corp., Littleton, CO USA 
Precision Navigation for a Mars Airplane 

Lowrie, James W.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 2, 196-197; In English; No 
Copyright; Avail: CASI; A01, Hardcopy 

The rough Martian terrain significantly impedes high speed travel by wheeled vehicles and much of it is simply 
inaccessible given the capability of typical rover designs. Airplanes, however, have much greater range and can provide access 
to scientifically interesting terrain that is inaccessible to landers and rovers. Moreover, they can provide coverage of a large 
portion of the surface and return high resolution images and science data not practical from orbiting spacecraft. Precise 
navigation on Earth requires a constellation of satellites such as GPS (Global Positioning Satellites) or a network of precisely 
located and calibrated ground beacons, an approach that is impractical for Mars exploration in the near future. In order to 
realize the benefits of airplane exploration on Mars, a precision navigation system is required. Such a system also provides 
a high degree of autonomous capability because it enables: (1) Accurate overflight of specifically targeted sites. (2) Hazard 
avoidance in low altitude flight. (3) The collection of 'focused' science data which reduces overall data volume and supports 
an optimized data return strategy (4) Accurate spatial and temporal correlation of acquired science data with orbiter 
observations. (5) A geodetically referenced site survey capability. (6) A soft landing capability by providing in-flight landing 
site selection and terminal guidance. (7) Return to a base station following flight. (8) Precise placement of science probes and 
future navigation beacons. SAIC's Center for Intelligent Systems (SAIC-CIS) leverages on experience from unmanned vehicle 
research to propose a concept for an intelligent landmark navigation system that relies on autonomous real-time recognition 
of visible surface features during flight. 
Author 
Mars Missions; Mars Exploration; Mars (Planet); Autonomous Navigation; Aircraft 

20010019289 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
MOLA-Based Landing Site Characterization 

Duxbury, T. C; Ivanov, A. B.; First Landing Site Workshop for the 2003 Mars Exploration Rovers; 2001, 18-19; In English; 
No Copyright; Abstract Only; Available from CASI only as part of the entire parent document 

The Mars Global Surveyor (MGS) Mars Orbiter Laser Altimeter (MOLA) data provide the basis for site characterization 
and selection never before possible. The basic MOLA information includes absolute radii, elevation and 1 micrometer albedo 
with derived datasets including digital image models (DIM's illuminated elevation data), slopes maps and slope statistics and 
small scale surface roughness maps and statistics. These quantities are useful in downsizing potential sites from descent 
engineering constraints and landing/roving hazard and mobility assessments. Slope baselines at the few hundred meter level 
and surface roughness at the 10 meter level are possible. Additionally, the MOLA-derived Mars surface offers the possibility 
to precisely register and map project other instrument datasets (images, ultraviolet, infrared, radar, etc.) taken at different 
resolution, viewing and lighting geometry, building multiple layers of an information cube for site characterization and 
selection. Examples of direct MOLA data, data derived from MOLA and other instruments data registered to MOLA arc given 
for the Hematite area. 
Author 
Mars Global Surveyor; Mars Surface; Landing Sites; Laser Altimeters 

20010002491 California Univ., Los Angeles, CA USA 
After the Mars Polar Lander: Where to Next? 

Paige, D. A.; International Conference on Mars Polar Science and Exploration; August 2000, 140-141; In English; No 
Copyright; Avail: CASI; A01, Hardcopy 

The recent loss of the Mars Polar Lander (MPL) mission represents a serious setback to Mars science and exploration. 
Targeted to land on the Martian south polar layered deposits at 76 deg south latitude and 195 deg west longitude, it would 
have been the first mission to study the geology, atmospheric environment, and volatiles at a high-latitude landing site. Since 
the conception of the MPL mission, a Mars exploration strategy has emerged which focuses on Climate, Resources and Life, 
with the behavior and history of water as the unifying theme. A successful MPL mission would have made significant 

27 



contributions towards these goals, particularly in understanding the distribution and behavior of near-surface water, and the 
nature and climate history of the south polar layered deposits. Unfortunately, due to concerns regarding the design of the MPL 
spacecraft, the rarity of direct trajectories that enable high-latitude landings, and funding, an exact reflight of MPL is not 
feasible within the present planning horizon. However, there remains significant interest in recapturing the scientific goals of 
the MPL mission. The following is a discussion of scientific and strategic issues relevant to planning the next polar lander 
mission, and beyond. Additional information is contained in the original extended abstract. 
Author 
Mars (Planet); Mars Exploration; Mars Polar Lander; Polar Regions; Mars Missions 

20000102372 Tennessee Univ., Knoxville, TN USA 

Earth Return Aerocapture for the TransHab/Ellipsled Vehicle 

Muth, W. D.; Hoffmann, G; Lyne, J. E.; October 2000; In English 
Contract(s)/Grant(s): NAG 1-2 163; No Copyright; Avail: CASI; A04, Hardcopy 

The current architecture being considered by NASA for a human Mars mission involves the use of an aerocapture 
procedure at Mars arrival and possibly upon Earth return. This technique would be used to decelerate the vehicles and insert 
them into their desired target orbits, thereby eliminating the need for propulsive orbital insertions. The crew may make the 
interplanetary journey in a large, inflatable habitat known as the TransHab. It has been proposed that upon Earth return, this 
habitat be captured into orbit for use on subsequent missions. In this case, the TransHab would be complimented with an 
aeroshell, which would protect it from heating during the atmospheric entry and provide the vehicle with aerodynamic lift. 
The aeroshell has been dubbed the 'Ellipsled' because of its characteristic shape. This paper reports the results of a preliminary 
study of the aerocapture of the TransHab/Ellipsled vehicle upon Earth return. Undershoot and overshoot boundaries have been 
determined for a range of entry velocities, and the effects of variations in the atmospheric density profile, the vehicle 
deceleration limit, the maximum vehicle roll rate, the target orbit, and the vehicle ballistic coefficient have been examined. A 
simple, 1 80 degree roll maneuver was implemented in the undershoot trajectories to target the desired 407 km circular Earth 
orbit. A three-roll sequence was developed to target not only a specific orbital energy, but also a particular inclination, thereby 
decreasing propulsive inclination changes and post-aerocapture delta-V requirements. Results show that the TransHab/ 
Ellipsled vehicle has a nominal corridor width of at least 0.7 degrees for entry speeds up to 14.0 km/s. Most trajectories were 
simulated using continuum flow aerodynamics, but the impact of high-altitude viscous effects was evaluated and found to be 
minimal. In addition, entry corridor comparisons have been made between the TransHab/Ellipsled and a modified Apollo 
capsule which is also being considered as the crew return vehicle; because of its slightly higher lift-to-drag ratio, the TransHab 
has a modest advantage with regard to corridor width. Stagnation-point heating rates and integrated heat loads were 
determined for a range of vehicle ballistic coefficients and entry velocities. 
Author 

Spacecraft Design; Product Development; Aerocapture; Aeromaneuvering; Interplanetary Transfer Orbits; Atmospheric 
Entry 

20000085950 North Carolina State Univ., Raleigh, NC USA 

An Investigation of Terminal Guidance and Control Techniques for a Robotic Mars Lander 

Birge, Brian K.; Walberg, Gerald; [2000]; In English 

Contract(s)/Grant(s): NAG1-2222; No Copyright; Avail: CASI; A03, Hardcopy 

Continuing on previous work, various precision landing control algorithms arc examined with the goal of minimizing the 
landed distance to a specified location on the Mars surface. This study considers a set of points from parachute handoff to 
touchdown on the surface. The first scenario considers a reverse gravity turn to a hover condition 500 meters above the surface 
and then uses lateral thrusting to minimize die range to target. The second scenario examines a guided, lifting parachute 
followed by a powered gravity turn to the targeted landing site. The third scenario considers thrust vectoring while on the 
ballistic parachute, followed by a reverse gravity turn to touchdown. 
Author 
Terminal Guidance; Command Guidance; Control Systems Design; Thrust Vector Control; Control Theory 

20000074639 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 

The Deep Space 4/Champollion Comet Rendezvous and Lander Technology Demonstration Mission 

Smythe, William D.; Weissman, Paul R.; Muirhead, Brian K.; Tan-Wang, Grace H.; Sabahi, Dara; Grimes, James M.; [2000]; 
In English; No Copyright; Avail: Other Sources; Abstract Only 

28 



The Deep Space 4/Champollion mission is designed to test and validate technologies for landing on and anchoring to 
small bodies, and sample collection and transfer, in preparation for future sample return missions from comets, asteroids, and 
satellites, in addition, DS-4 will test technologies for advanced, multi-engine solar electric propulsion (SEP) systems, 
inflatable-rigidizable solar arrays, autonomous navigation and precision guidance for landing, autonomous hazard detection 
and avoidance, and advanced integrated avionics and packaging concepts. Deep Space-4/Champollion consists of two 
spacecraft: an orbiter/carrier vehicle which includes the multi-engine SEP stage, and a lander, called Champollion, which will 
descend to the surface of the 46P/Tempel 1 cometary nucleus. The spacecraft will launch in April, 2003 and land on the comet 
in September, 2006 Deep Space 4/Champollion is a joint project between NASA and CNES, the French space agency. 
Author 
Deep Space; Space Missions; Mission Planning; Landing; Comets; Asteroids 



20000074247 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
Navigation Strategy for the Mars 2001 Lander Mission 

Mase, Robert A.; Spencer, David A.; Smith, John C; Braun, Robert D.; [2000]; In English; No Copyright; Avail: Other 
Sources; Abstract Only 

The Mars Surveyor Program (MSP) is an ongoing series of missions designed to robotically study, map and search for 
signs of life on the planet Mars. The MSP 2001 project will advance the effort by sending an orbiter, a lander and a rover to 
the red planet in the 2001 opportunity. Each vehicle will carry a science payload that will Investigate the Martian environment 
on both a global and on a local scale. Although this mission will not directly search for signs of life, or cache samples to be 
returned to Earth, it will demonstrate certain enabling technologies that will be utilized by the future Mars Sample Return 
missions. One technology that is needed for the Sample Return mission is the capability to place a vehicle on the surface 
within several kilometers of the targeted landing site. The MSP'01 Lander will take the first major step towards this type of 
precision landing at Mars. Significant reduction of the landed footprint will be achieved through two technology advances. The 
first, and most dramatic, is hypersonic aeromaneuvering; the second is improved approach navigation. As a result, the guided 
entry will produce in a footprint that is only tens of kilometers, which is an order of magnitude improvement over the 
Pathfinder and Mars Polar Lander ballistic entries. This reduction will significantly enhance scientific return by enabling the 
potential selection of otherwise unreachable landing sites with unique geologic interest and public appeal. A landed footprint 
reduction from hundreds to tens of kilometers is also a milestone on the path towards human exploration of Mars, where the 
desire is to place multiple vehicles within several hundred meters of the planned landing site. Hypersonic aeromaneuvering 
is an extension of the atmospheric flight goals of the previous landed missions, Pathfinder and Mars Polar Lander (MPL), that 
utilizes aerodynamic lift and an autonomous guidance algorithm while in the upper atmosphere. The onboard guidance 
algorithm will control the direction of the lift vector, via bank angle modulation, to keep the vehicle on the desired trajectory. 
While numerous autonomous guidance algorithms have been developed for use during hypersonic flight at Earth, this will be 
the first flight of an autonomously directed lifting entry vehicle at Mars. However, without sufficient control and knowledge 
of the atmospheric entry conditions, the guidance algorithm will not perform effectively. The goal of the interplanetary 
navigation strategy is to deliver the spacecraft to the desired entry condition with sufficient accuracy and knowledge to enable 
satisfactory guidance algorithm performance. Specifically, the entry flight path angle must not exceed 0.27 deg. to a 3 sigma 
confidence level. Entry errors will contribute directly to the size of the landed footprint and the most significant component 
is entry flight path angle. The size of the entry corridor is limited on the shallow side by integrated heating constraints, and 
on the steep side by deceleration (g-load) and terminal descent propellant. In order to meet this tight constraint it is necessary 
to place a targeting maneuver seven hours prior to the time of entry. At this time the trajectory knowledge will be quite 
accurate, and the effects of maneuver execution errors will be small. The drawback is that entry accuracy is dependent on the 
success of this final late maneuver. Because propulsive maneuvers are critical events, it is desirable to minimize their 
occurrence and provide the flight team with as much response time as possible in the event of a spacecraft fault. A mission 
critical maneuver at Entry - 7 hours does not provide much fault tolerance, and it is desirable to provide a strategy that 
minimizes reliance on this maneuver. This paper will focus on the Improvements in interplanetary navigation that will decrease 
entry errors and will reduce the landed footprint, even in the absence of aeromaneuvering. The easiest to take advantage of 
are Improvements In the knowledge of the Mars ephemeris and gravity field due to the MGS and MSP'98 missions. 
Improvements In data collection and reduction techniques such as 'precision ranging' and near-simultaneous tracking will also 
be utilized. In addition to precise trajectory control, a robust strategy for communications and flight operations must also be 
demonstrated. The result Is a navigation and communications strategy on approach that utilizes optimal maneuver placement 
to take advantage of trajectory knowledge, minimizes risk for the flight operations team, is responsive to spacecraft hardware 
limitations, and achieves the entry corridor. The MSP2001 mission Is managed at JPL under the auspices of the Mars 

29 



Exploration Directorate. The spacecraft flight elements are built and managed by Lockheed-Martin Astronautics in Denver, 

Colorado. 

Author 

Interplanetary Navigation; Landing Sites; Mars Landing; Earth-Mars Trajectories; Orbital Mechanics; Orbit Calculation; 

Mars Surveyor 2001 Mission 

20000074083 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
A Light- Weight Inflatable Hypersonic Drag Device for Planetary Entry 

McRonald, Angus D.; [2000]; In English; No Copyright; Avail: Other Sources; Abstract Only 

The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, for flight in planetary 
atmospheres, for entry, aerocapture, and aerobraking. Studies to date include Mars, Venus, Earth, Saturn, Titan, Neptune and 
Pluto, and data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using 
a ballute is that aero, deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute 
than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry 
speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the heating 
by a factor of 10 for the spacecraft and a factor of 30 for the ballute. Consequently the entry pay load (lander, orbiter, etc) is 
subject to much less heating, requires a much reduced thermal, protection system (possibly only an MLI blanket), and the 
spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough 
to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more 
shallow angle, even allowing for the increased delivery angle error. Added advantages are less mass ratio of entry system to 
total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies 
suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an 
orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint. 
The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute 
adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a 
lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large 
potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute 
or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively 
small entry speeds, such as for Mars and Titan, or for the extensive atmosphere of a low-gravity planet such as Pluto. Details 
of a ballute to place a small Mars orbiter and a small Pluto lander will be given to illustrate the concept. The author will discuss 
presently available ballute materials and a development program of aerodynamic tests and materials that would be required 
for ballutes to achieve their full potential. 
Author 

Aerodynamic Heating; Research; Ballutes; Buoyancy; Drag Devices; Floats; Inflatable Structures; Microgravity; Planetary 
Atmospheres; Spacecraft Design 

20000062309 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
Precise Image-Based Motion Estimation for Autonomous Small Body Exploration 

Johnson, Andrew Edie; Matthies, Larry H.; [2000]; In English; 5th; No Copyright; Avail: CASI; A01, Hardcopy 

We have developed and tested a software algorithm that enables onboard autonomous motion estimation near small bodies 
using descent camera imagery and laser altimetry. Through simulation and testing, we have shown that visual feature tracking 
can decrease uncertainty in spacecraft motion to a level that makes landing on small, irregularly shaped, bodies feasible. 
Possible future work will include qualification of the algorithm as a flight experiment for the Deep Space 4/Champollion comet 
lander mission currently under study at the Jet Propulsion Laboratory. 
Author 
Estimating; Autonomy; Spacecraft Motion; Optical Tracking; Image Analysis; Algorithms 

20000057306 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 

Aerobraking at Venus and Mars: A Comparison of the Magellan and Mars Global Surveyor Aerobraking Phases 

Lyons, Daniel T; [2000]; In English; No Copyright; Avail: CASI; A01, Hardcopy 

On February 4, 1999 the Mars Global Surveyor spacecraft became the second spacecraft to successfully aerobrake into 
a nearly circular orbit about another planet. This paper will highlight some of the similarities and differences between the 
aerobraking phases of this mission and the first mission to use aerobraking, the Magellan mission to Venus. Although the Mars 

30 



Global Surveyor (MGS) spacecraft was designed for aerobraking and the Magellan spacecraft was not, aerobraking MGS was 
a much more challenging task than aerobraking Magellan, primarily because the spacecraft was damaged during the initial 
deployment of the solar panels. The MGS aerobraking phase had to be completely redesigned to minimize the bending 
moment acting on a broken yoke connecting one of the solar panels to the spacecraft. Even if the MGS spacecraft was 
undamaged, aerobraking at Mars was more challenging than aerobraking at Venus for several reasons. First, Mars is subject 
to dust storms, which can significantly change the temperature of the atmosphere due to increased solar heating in the low and 
middle altitudes (below 50 km), which in turn can significantly increase the density at the aerobraking altitudes (above 100 
km). During the first part of the MGS aerobraking phase, a regional dust storm was observed to have a significant and very 
rapid effect on the entire atmosphere of Mars. Computer simulations of global dust storms on Mars indicate that even larger 
density increases are possible than those observed during the MGS aerobraking phases. For many aerobraking missions, the 
duration of the aerobraking phase must be kept as short as possible to minimize the total mission cost. For Mars missions, a 
short aerobraking phase means that there will be less margin to accommodate atmospheric variability, so the operations team 
must be ready to propulsively raise periapsis by tens of kilometers on very short notice. This issue was less of a concern on 
Venus, where the thick lower atmosphere and the slow planet rotation resulted in more predictable atmospheric densities from 
one orbit to the next. 
Author 
Aerobraking; Mars Global Surveyor; Magellan Spacecraft (NASA); Circular Orbits 



20000056881 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA 
The Strategy for the Second Phase of Aerobraking Mars Global Surveyor 

Johnston, M. D.; Esposito, P. B.; Alwar, V; Demcak, S. W.; Graat, E. J.; Burkhart, P. D.; Portock, B. M.; [2000]; In English; 
No Copyright; Avail: CASI; A01, Hardcopy 

On February 19, 1999, the Mars Global Surveyor (MGS) spacecraft was able to propulsively establish its mapping orbit. 
This event followed the completion of the second phase of aerobraking for the MGS spacecraft on February 4, 1999. For the 
first time, a spacecraft at Mars had successfully employed aerobraking methods in order to reach its desired pre-launch 
mapping orbit. This was accomplished despite a damaged spacecraft solar array. The MGS spacecraft was launched on 
November 7, 1996, and after a ten month interplanetary transit was inserted into a highly elliptical capture orbit at Mars on 
September 12, 1997. Unlike other interplanetary missions, the MGS spacecraft was launched with a planned mission delta-V 
((Delta) V) deficit of nearly 1250 m/s. To overcome this AV deficit, aerobraking techniques were employed. However, damage 
discovered to one of the spacecraft's two solar arrays after launch forced major revisions to the original aerobraking planning 
of the MGS mission. In order to avoid a complete structural failure of the array, peak dynamic pressure levels for the spacecraft 
were established at a major spacecraft health review in November 1997. These peak dynamic pressure levels were roughly 
one-third of the original mission design values. Incorporating the new dynamic pressure limitations into mission replanning 
efforts resulted in an 'extended' orbit insertion phase for the mission. This 'extended' orbit insertion phase was characterized 
by two distinct periods of aerobraking separated by an aerobraking hiatus that would last for several months in an intermediate 
orbit called the 'Science Phasing Orbit' (SPO). This paper describes and focuses on the strategy for the second phase of 
aerobraking for the MGS mission called 'Aerobraking Phase 2.' This description will include the baseline aerobraking flight 
profile, the trajectory control methodology, as well as the key trajectory metrics that were monitored in order to successfully 
'guide' the spacecraft to its desired mapping orbit. Additionally, the actual aerobraking progress is contrasted to the planned 
aerobraking flight profile. (A separate paper will describe the navigation aspects of MGS aerobraking in detail.) Key to the 
success of the MGS mission is the delivery of the spacecraft to its final mapping orbit and the synergy the instrument 
complement provides to its scientific investigators when science data is returned from that orbit. The MGS mapping orbit is 
characterized as a low altitude, near-circular, near-polar orbit that is Sun-synchronous with the descending equatorial crossing 
at 2:00 AM local mean solar time (LMST). 
Derived from text 
Aerobraking; Mars Global Surveyor; Mapping; Orbit Insertion; Trajectory Control; Navigation; Mission Planning 



31 



Subject Terms 



2001 MARS ODYSSEY 

Application of Accel era meter Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

ACCELEROMETERS 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Approaches to autonomous aerobraking 
at Mars - 1 5 

ACOUSTO-OPTICS 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 

ACTIVE CONTROL 

Precision Terminal Guidance for a Mars 
Lander - 25 

AEROASSIST 

Aeroassist Technology Planning for 
Exploration - 4 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

NASA Development of Aerocapture 
Technologies - 14 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

AEROBRAKING 

Aerobraking at Venus and Mars: A Com- 
parison of the Magellan and Mars Global 
Surveyor Aerobraking Phases - 30 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Autonomous Aerobraking at Mars - 16 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 



AEROCAPTURE 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Aerocapture Technology Project 
Overview - 11 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 

- 9 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Atmospheric Models for Aerocapture - 4 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

Mars Sample Return without Landing on 
the Surface - 25 

NASA Development of Aerocapture 
Technologies - 14 

Neptune Aerocapture Systems Analysis 

- 3 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

AERODYNAMIC CHARACTERISTICS 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

AERODYNAMIC DRAG 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

AERODYNAMIC FORCES 

NASA Development of Aerocapture 
Technologies - 14 



AERODYNAMIC HEATING 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

AERODYNAMIC LOADS 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

AERODYNAMICS 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 
- 12 



AEROMANEUVERING 

Autonomous Aerobraking at Mars 



16 



Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

AEROSHELLS 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

Neptune Aerocapture Systems Analysis 

- 3 

AEROSPACE SCIENCES 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

Daily repeat-groundtrack Mars orbits - 7 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

Europa Lander - 13 

Pioneer Venus and Galileo entry probe 
heritage - 8 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 



A-1 



AEROTHERMODYNAMICS 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

AIR BAG RESTRAINT DEVICES 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

AIRCRAFT 

Precision Navigation for a Mars Airplane 

- 27 

ALGORITHMS 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Aerocapture Guidance Methods for High 
Energy Trajectories - 11 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 

- 9 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

ANGLE OF ATTACK 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 

- 9 

ANNUAL VARIATIONS 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

APPLICATIONS PROGRAMS (COMPUT- 
ERS) 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

ASTEROIDS 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

ASTROPHYSICS 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

ATMOSPHERIC CHEMISTRY 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

ATMOSPHERIC DENSITY 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 



Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

ATMOSPHERIC ENTRY 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Computational Analysis of Towed Ballute 
Interactions - 18 

Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 1 4 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 16 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

ATMOSPHERIC MODELS 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Atmospheric Models for Aerocapture - 4 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

ATMOSPHERIC PHYSICS 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

AUTONOMOUS NAVIGATION 

Micro Navigator - 23 

Precision Navigation for a Mars Airplane 

- 27 

AUTONOMY 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

BALLUTES 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Computational Analysis of Towed Ballute 
Interactions - 18 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

BUOYANCY 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

CAMERAS 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 



Optical landmark detection for spacecraft 
navigation - 6 

CAPTURE EFFECT 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

CIRCULAR ORBITS 

Aerobraking at Venus and Mars: A Com- 
parison of the Magellan and Mars Global 
Surveyor Aerobraking Phases - 30 

COLLISIONS 

Low Velocity Impact Experiments in 
Microgravity - 24 

COMETS 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

COMMAND GUIDANCE 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 

COMMUNICATION SATELLITES 

Daily repeat-groundtrack Mars orbits - 7 

COMPUTATIONAL FLUID DYNAMICS 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

COMPUTERIZED SIMULATION 

Autonomous Aerobraking at Mars - 16 

Computational Analysis of Towed Ballute 
Interactions - 18 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 

- 9 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 16 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 

- 12 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

COMPUTERS 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 6 

CONTROL SURFACES 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

CONTROL SYSTEMS DESIGN 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 



A-2 



CONTROL THEORY 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 

- 9 

CONVECTIVE HEAT TRANSFER 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

CORRELATION 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

DATA ACQUISITION 

Approach navigation for the 2009 Mars 
large lander - 6 

DATA BASES 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

DEATH VALLEY (CA) 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

DEEP SPACE 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

DEGREES OF FREEDOM 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 14 

DESCENT 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 

- 9 

DIFFERENTIAL EQUATIONS 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 16 

DRAG DEVICES 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

DRAG 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 

- 12 

DUST 

Low Velocity Impact Experiments in 
Microgravity - 24 

EARTH-MARS TRAJECTORIES 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

ELLIPTICAL ORBITS 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 



ENERGY DISSIPATION 

Low Velocity Impact Experiments in 
Microgravity - 24 

ENVIRONMENT PROTECTION 

International Agreement on Planetary 
Protection - 24 

ENVIRONMENTAL MONITORING 

Atmospheric Models for Aerocapture - 4 

ESTIMATING 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 
- 30 

EUROPEAN SPACE PROGRAMS 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

EVALUATION 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

EXOBIOLOGY 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

FEASIBILITY ANALYSIS 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

FLIGHT CONTROL 

Computational Analysis of Towed Ballute 
Interactions - 18 



FLIGHT OPERATIONS 

Autonomous Aerobraking at Mars 



16 



The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

FLIGHT SIMULATION 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 16 

FLOATS 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

FLOW CHARACTERISTICS 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

FLOW DISTRIBUTION 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

GALILEAN SATELLITES 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

GALILEO SPACECRAFT 

Pioneer Venus and Galileo entry probe 
heritage - 8 

GAS GIANT PLANETS 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 



GRAND TOURS 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

GRAVITATIONAL EFFECTS 

Low Velocity Impact Experiments in 
Microgravity - 24 

GROUND TRACKS 

Daily repeat-groundtrack Mars orbits - 7 

GUIDANCE (MOTION) 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 



HEAT FLUX 

Autonomous Aerobraking at Mars 



16 



HEAT SHIELDING 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 

HEAT TRANSFER COEFFICIENTS 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

HIGH RESOLUTION 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

HUYGENS PROBE 

Atmospheric Models for Aerocapture - 4 

HYPERBOLIC TRAJECTORIES 

Aerocapture Guidance Methods for High 
Energy Trajectories - 11 

HYPERSONIC FLOW 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

HYPERSONIC WAKES 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

IMAGE ANALYSIS 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

IMAGING TECHNIQUES 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

IMPACT 

Low Velocity Impact Experiments in 
Microgravity - 24 

IMPINGEMENT 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

IN SITU MEASUREMENT 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 



A-3 



INFLATABLE STRUCTURES 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

INTERNATIONAL COOPERATION 

International Agreement on Planetary 
Protection - 24 

INTERNATIONAL LAW 

International Agreement on Planetary 
Protection - 24 

INTERNATIONAL SPACE STATION 

Aeroassist Technology Planning for 
Exploration - 4 

INTERORBITAL TRAJECTORIES 

Optimization of Low Thrust Trajectories 
With Terminal Aerocapture - 11 

INTERPLANETARY NAVIGATION 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

INTERPLANETARY SPACECRAFT 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

INTERPLANETARY TRAJECTORIES 

Optimization of Low Thrust Trajectories 
With Terminal Aerocapture - 11 

INTERPLANETARY TRANSFER ORBITS 

Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

INVISCID FLOW 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

ION PROPULSION 

SEP Mission to Titan NEXT Aerocapture 
In-Space Propulsion (Quicktime Movie) 
- 10 

KALMAN FILTERS 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

LANDING MODULES 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 



LASER ALTIMETERS 

MOLA-Based Landing 

Characterization - 27 



Site 



LANDING SITES 

MOLA-Based 
Characterization 



Landing 
27 



Site 



Navigation Strategy for the Mars 2001 
Lander Mission - 29 

LANDING 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

LAND 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 



LOADS (FORCES) 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

LOW SPEED 

Low Velocity Impact Experiments in 
Microgravity - 24 

MAGELLAN SPACECRAFT (NASA) 

Aerobraking at Venus and Mars: A Com- 
parison of the Magellan and Mars Global 
Surveyor Aerobraking Phases - 30 

MANNED MARS MISSIONS 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

MAPPING 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

MARS ATMOSPHERE 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

MARS EXPLORATION 

After the Mars Polar Lander: Where to 
Next? - 26 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 1 4 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 

- 9 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Precision Navigation for a Mars Airplane 

- 27 

The Martian Oasis Detector - 26 

MARS GLOBAL SURVEYOR 

Aerobraking at Venus and Mars: A Com- 
parison of the Magellan and Mars Global 
Surveyor Aerobraking Phases- 30 



MOLA-Based 
Characterization 



Landing 
27 



Site 



The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

MARS LANDING 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

Precision Terminal Guidance for a Mars 
Lander - 25 



MARS MISSIONS 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

After the Mars Polar Lander: Where to 
Next? - 26 

Autonomous Aerobraking at Mars - 16 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Lunar and Planetary Science XXXV: Mis- 
sions and Instruments: Hopes and Hope 
Fulfilled - 5 

Precision Navigation for a Mars Airplane 

- 27 

The Martian Oasis Detector - 26 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

MARS (PLANET) 

After the Mars Polar Lander: Where to 
Next? - 26 

Precision Navigation for a Mars Airplane 

- 27 

The Martian Oasis Detector - 26 

MARS POLAR LANDER 

After the Mars Polar Lander: Where to 
Next? - 27 

MARS ROVING VEHICLES 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 1 4 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 

- 9 

MARS SAMPLE RETURN MISSIONS 

Mars Sample Return without Landing on 
the Surface - 25 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 



MARS SURFACE 

MOLA-Based 
Characterization 

The Martian Oasis Detector 



Landing 
27 



Site 



26 



MARS SURVEYOR 2001 MISSION 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

MATHEMATICAL MODELS 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 
- 9 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

MEASURING INSTRUMENTS 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

MECHANICAL OSCILLATORS 

Ultra-stable oscillators for planetary entry 
probes - 8 



A-4 



METEORITES 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

MICROGRAVITY 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Low Velocity Impact Experiments in 
Microgravity - 24 

MINERAL DEPOSITS 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 



MINIATURIZATION 

Micro Navigator - 



23 



MISSION PLANNING 

After the Mars Polar Lander: Where to 
Next? - 26 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

MONTE CARLO METHOD 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

MULTISENSOR FUSION 

Micro Navigator - 23 

NASA SPACE PROGRAMS 

Aeroassist Technology Planning for 
Exploration - 4 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 14 

NASA Development of Aerocapture 
Technologies - 14 

NAVIGATION 

Approach navigation for the 2009 Mars 
large lander - 6 

Approaches to autonomous aerobraking 
at Mars - 1 5 

Optical landmark detection for spacecraft 
navigation - 6 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

NEPTUNE ATMOSPHERE 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 



NEPTUNE (PLANET) 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 

- 9 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

Neptune Aerocapture Systems Analysis 

- 3 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

NUCLEAR ELECTRIC PROPULSION 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

NUMERICAL ANALYSIS 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

OASES 

The Martian Oasis Detector - 26 

OPTICAL TRACKING 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

OPTIMAL CONTROL 

Optimization of Low Thrust Trajectories 
With Terminal Aerocapture - 11 

ORBIT CALCULATION 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

ORBIT INSERTION 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

ORBITAL MECHANICS 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

ORBITS 

Daily repeat-groundtrack Mars orbits - 7 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

OUTER PLANETS EXPLORERS 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

PARACHUTES 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 6 

PARTICULATE SAMPLING 

Mars Sample Return without Landing on 
the Surface - 25 



PERIODIC VARIATIONS 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

PLANETARY ATMOSPHERES 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

PLANETARY ENVIRONMENTS 

International Agreement on Planetary 
Protection - 24 

PLANETARY LANDING 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 

Approach navigation for the 2009 Mars 
large lander - 6 

Daily repeat-groundtrack Mars orbits - 7 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

Europa Lander - 13 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 

PLANETOLOGY 

Lunar and Planetary Science XXXV: Mis- 
sions and Instruments: Hopes and Hope 
Fulfilled - 5 

PLANETS 

Approaches to autonomous aerobraking 
at Mars - 1 5 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

Daily repeat-groundtrack Mars orbits - 7 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

Pioneer Venus and Galileo entry probe 
heritage - 8 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 



A-5 



PLANNING 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

POLAR REGIONS 

After the Mars Polar Lander: Where to 
Next? - 26 

POLICIES 

International Agreement on Planetary 
Protection - 24 

POSITION (LOCATION) 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 

POWERED LIFT AIRCRAFT 

Exploration of Titan Using Vertical Lift 
Aerial Vehicles - 23 

PREDICTIONS 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

PREDICTOR-CORRECTOR METHODS 

Aerocapture Guidance Methods for High 
Energy Trajectories - 11 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

PRESSURE VESSELS 

Pioneer Venus and Galileo entry probe 
heritage - 8 

PRODUCT DEVELOPMENT 

Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

PROPULSION 

NASA Development of Aerocapture 
Technologies - 14 

QUARTZ 

Ultra-stable oscillators for planetary entry 
probes - 8 

RADIATIVE HEAT TRANSFER 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

RADIO TRACKING 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

RADIOACTIVE ISOTOPES 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

RECONNAISSANCE AIRCRAFT 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

REMOTE SENSING 

Remote Sensing of Evaporite Minerals in 
Bad water Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

RESEARCH 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 



RESOLUTION 

Optical landmark detection for spacecraft 
navigation - 6 

ROBOTICS 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

ROCKET EXHAUST 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

ROTARY WING AIRCRAFT 

Exploration of Titan Using Vertical Lift 
Aerial Vehicles - 23 

ROVING VEHICLES 

Approach navigation for the 2009 Mars 
large lander - 6 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

RUBIDIUM 

Ultra-stable oscillators for planetary entry 
probes - 8 

SATELLITE SURFACES 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

SEDIMENTARY ROCKS 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

SIMULATION 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 



SOLAR ARRAYS 

Autonomous Aerobraking at Mars 



16 



Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

SOLAR ELECTRIC PROPULSION 

SEP Mission to Titan NEXT Aerocapture 
In-Space Propulsion (Quicktime Movie) 
- 10 

SOLAR SYSTEM 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Europa Lander - 13 

SPACE EXPLORATION 

Aeroassist Technology Planning for 
Exploration - 4 



Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Exploration of Titan Using Vertical Lift 
Aerial Vehicles - 23 

SPACE FLIGHT 

Approaches to autonomous aerobraking 
at Mars - 1 5 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

SPACE MISSIONS 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 
- 9 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

SPACE NAVIGATION 

Micro Navigator - 23 

SPACE PROBES 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Pioneer Venus and Galileo entry probe 
heritage - 8 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 

SPACECRAFT CONTROL 

Computational Analysis of Towed Ballute 
Interactions - 18 

SPACECRAFT DESIGN 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

SPACECRAFT GUIDANCE 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Micro Navigator - 23 

Precision Terminal Guidance for a Mars 
Lander - 25 

SPACECRAFT INSTRUMENTS 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 



A-6 



Lunar and Planetary Science XXXV: Mis- 
sions and Instruments: Hopes and Hope 
Fulfilled - 5 

The Martian Oasis Detector - 26 

SPACECRAFT MANEUVERS 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Autonomous Aerobraking at Mars - 16 

SPACECRAFT MOTION 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

SPACECRAFT PERFORMANCE 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

SPACECRAFT PROPULSION 

Aerocapture Technology Project 
Overview - 11 

Optical landmark detection for spacecraft 
navigation - 6 

SPACECRAFT 

Approaches to autonomous aerobraking 
at Mars - 1 5 

Europa Lander - 13 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 6 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 

- 12 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

SPATIAL RESOLUTION 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

SPECTRAL BANDS 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

STATISTICAL ANALYSIS 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

STIFFNESS 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 6 

STRUCTURAL BASINS 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

STRUCTURAL DESIGN 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

SURFACE PROPERTIES 

Science and Engineering Potential of an 
Icy Moon Lander - 13 



SYSTEMS ANALYSIS 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Aerocapture Technology Project 
Overview - 11 

NASA Development of Aerocapture 
Technologies - 14 

Neptune Aerocapture Systems Analysis 

- 3 

TECHNOLOGICAL FORECASTING 

Aeroassist Technology Planning for 
Exploration - 4 

TECHNOLOGY UTILIZATION 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Neptune Aerocapture Systems Analysis 

- 3 

TEMPERATURE MEASURING INSTRU- 
MENTS 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 



TEMPERATURE PROFILES 

Autonomous Aerobraking at Mars 



16 



TERMINAL GUIDANCE 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 
- 9 

Precision Terminal Guidance for a Mars 
Lander - 25 

THERMAL ANALYSIS 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

THRUST VECTOR CONTROL 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 

TITAN 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

TOPOGRAPHY 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

TOWED BODIES 

Computational Analysis of Towed Ballute 
Interactions - 18 

TRACKING (POSITION) 

Optical landmark detection for spacecraft 
navigation - 6 



TRAJECTORIES 

Aerocapture Guidance Methods for High 
Energy Trajectories - 11 

Approach navigation for the 2009 Mars 
large lander - 6 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 6 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

TRAJECTORY ANALYSIS 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 1 4 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 
- 9 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

TRAJECTORY CONTROL 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

TRAJECTORY OPTIMIZATION 

Optimization of Low Thrust Trajectories 
With Terminal Aerocapture - 11 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

TRAJECTORY PLANNING 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

TRANSFER ORBITS 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

TRANSPONDERS 

Ultra-stable oscillators for planetary entry 
probes - 8 

TRITON 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

VAPORIZING 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 



A-7 



VENUS (PLANET) WATER Control Surface and Afterbody Experi- 

Pioneer Venus and Galileo entry probe The Martian Oasis Detector - 26 mental Aeroheating for a Proposed Mars 

heritage- 8 Smart Lander Aeroshell - 17 

VERTICAL TAKEOFF AIRCRAFT WIND TUNNEL TESTS Wake aosure characteristics and After- 
Exploration of Titan Using Vertical Lift CFD Prediction of the BEAGLE 2 Mars body Heating on a Mars Sample Return 
Aerial Vehicles - 23 Probe Aerodynamic Database - 19 Orbiter- 15 



A-8 



Corporate Sources 



Air Force Inst, of Tech. 

Aerocapture Guidance Methods for High 
Energy Trajectories - 11 



Arizona Univ. 

The Martian Oasis Detector - 



26 



Ball Aerospace and Technologies Corp. 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

California Univ. 

After the Mars Polar Lander: Where to 
Next? - 26 

Centre National d'Etudes Spatiales 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

CFD Research Corp. 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Colorado Univ. 

Low Velocity Impact Experiments in 
Microgravity - 24 

Computer Sciences Corp. 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

Georgia Inst, of Tech. 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

Instituto Nacional de Pesquisas Espacias 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

Jet Propulsion Lab., California Inst, of 
Tech. 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Aerobraking at Venus and Mars: A Com- 
parison of the Magellan and Mars Global 
Surveyor Aerobraking Phases - 30 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Mars Sample Return without Landing on 
the Surface - 25 



Micro Navigator - 23 



MOLA-Based 
Characterization 



Landing 
27 



Site 



Navigation Strategy for the Mars 2001 
Lander Mission - 29 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 
- 30 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 



The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Lunar and Planetary Inst. 

Lunar and Planetary Science XXXV: Mis- 
sions and Instruments: Hopes and Hope 
Fulfilled - 5 

Morgan Research Corp. 

Atmospheric Models for Aerocapture - 4 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

NASA Ames Research Center 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Exploration of Titan Using Vertical Lift 
Aerial Vehicles - 23 

International Agreement on Planetary 
Protection - 24 

NASA Glenn Research Center 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

NASA Johnson Space Center 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

NASA Langley Research Center 

Aeroassist Technology Planning for 
Exploration - 4 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Autonomous Aerobraking at Mars - 16 

Computational Analysis of Towed Ballute 
Interactions - 18 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 1 4 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 
- 9 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 



Mars Smart Lander Parachute Simula- 
tion Model - 19 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 16 

Neptune Aerocapture Systems Analysis 

- 3 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

NASA Marshall Space Flight Center 

Aerocapture Technology Project 
Overview - 11 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

NASA Development of Aerocapture 
Technologies - 14 

SEP Mission to Titan NEXT Aerocapture 
In-Space Propulsion (Quicktime Movie) 

- 10 

Naval Postgraduate School 

Optimization of Low Thrust Trajectories 
With Terminal Aerocapture - 11 

North Carolina State Univ. 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 

Rhode Island Univ. 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

Science Applications International Corp. 

Precision Navigation for a Mars Airplane 

- 27 

Precision Terminal Guidance for a Mars 
Lander - 25 



Tennessee Univ. 

Earth Return Aerocapture for 
TransHab/Ellipsled Vehicle - 28 



the 



Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 



B-1 



Document Authors 



Alter, Stephen J. 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

Alwar, V. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Amundsen, Ruth M. 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Anderson, Brian P. 

Computational Analysis of Towed Ballute 
Interactions - 18 

Asmar, S. W. 

Ultra-stable oscillators for planetary entry 
probes - 8 

Atkinson, D. H. 

Ultra-stable oscillators for planetary entry 
probes - 8 

Atkinson, David 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Atreya, Sushil 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Bada, J. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Baggett, Randy 

SEP Mission to Titan NEXT Aerocapture 
In-Space Propulsion (Quicktime Movie) 

- 10 

Baldridge, A. 

Remote Sensing of Evaporite Minerals in 
Bad water Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

Banfield, Donald 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Beaty, D. W. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Beebe, Reta 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 



Benson, Scott 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

Bienstock, Bernard J. 

Pioneer Venus and Galileo entry probe 
heritage - 8 

Bird, M. K. 

Ultra-stable oscillators for planetary entry 
probes - 8 

Birge, Brian K. 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 



Blaes, B. R. 

Micro Navigator - 



23 



Blanchard, Robert C. 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

Blaney, Diana L. 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 

Bolton, Scott 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Boynton, W. V. 

After the Mars Polar Lander: Where to 
Next? - 26 

Braun, Robert D. 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

Briggs, Geoffrey 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Burkhart, P. Daniel 

Approach navigation for the 2009 Mars 
large lander - 6 

Burkhart, P. D. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Burnell, Simon I. 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Chapel, J. 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

Charbonnier, J. M. 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 



Chau, S. N. 

Micro Navigator - 



23 



Chavis, Zachary Q. 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

Cheatwood, F. M. 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 14 

Cheatwood, F. McNeil 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Cheatwood, McNeil F. 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

Cheng, Yang 

Optical landmark detection for spacecraft 
navigation - 6 

Cianciolo, Alicia Dwyer 

Autonomous Aerobraking at Mars - 16 

Cledassou, R. 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

Colwell, J. E. 

Low Velocity Impact Experiments in 
Microgravity - 24 

Conrad, P. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 

Crisp, D. 

After the Mars Polar Lander: Where to 
Next? - 26 

Crisp, David 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Cutts, James 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Cwynar, D. 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

Dec, John A. 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

Dec, John 

Autonomous Aerobraking at Mars - 16 

DeJong, E. 

After the Mars Polar Lander: Where to 
Next? - 26 



C-1 



Delamere, A. 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

Demcak, S. W. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Desai, Prasun N. 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 14 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

DHondt, S. L. 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

Dicarlo, Jennifer L. 

Aerocapture Guidance Methods for High 
Energy Trajectories - 11 

Diehl, Roger 

Daily repeat-groundtrack Mars orbits - 7 

Drake, Michael 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Dupuis, E. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 

Duvall, Aleta L. 

Atmospheric Models for Aerocapture - 4 

Duvall, Aleta 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 



Duxbury, T. C. 

MOLA-Based 
Characterization 



Landing 
27 



Site 



Duxbury, Thomas C. 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Dwyer, A. M. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Dwyer, Alicia M. 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 



Dyke, R. Eric 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

Edquist, Karl T. 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

Ely, Todd 

Daily repeat-groundtrack Mars orbits - 7 

Escalera, P. E. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Esposito, Larry 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Esposito, P. B. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

et al. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Farmer, J. 

Remote Sensing of Evaporite Minerals in 
Badwater Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 

Galal, Kenneth 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Gasbarre, Joseph F. 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 

Gefert, Leon 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

Gehling, R. 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

George, B. E. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

George, Benjamin E. 

Aeroheating Thermal Analysis Methods 
for Aerobraking Mars Missions - 4 

Thermal Analysis and Correlation of the 
Mars Odyssey Spacecraft's Solar Array 
During Aerobraking Operations - 17 



Gibson, Everett K., Jr. 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Glenar, David A. 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 

Gnoffo, Peter A. 

Computational Analysis of Towed Ballute 
Interactions - 18 

Golombek, Matthew P. 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Graat, E. J. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Graves, Claude 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Grimes, James M. 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

Gulick, Doug 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Habchi, Sami D. 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Hall, Jeff 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Hanna, J. L. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Approaches to autonomous aerobraking 
at Mars - 1 5 



Hanna, Jill L. 

Autonomous Aerobraking at Mars 



16 



Hansen, C. J. 

After the Mars Polar Lander: Where to 
Next? - 26 

Harri, A. M. 

After the Mars Polar Lander: Where to 
Next? - 26 

Hillman, John J. 

AIMS: Acousto-optic imaging spectrom- 
eter for spectral mapping of solid 
surfaces - 13 



Hoffmann, C. 

Earth Return Aerocapture for 
TransHab/Ellipsled Vehicle - 28 



the 



Gershman, Robert 

Europa Lander - 



13 



Hollis, Brian R. 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 



C-2 



Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Horvath, Thomas J. 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

Hrinda, Glenn A. 

Structural Design for a Neptune Aerocap- 
ture Mission - 2 

Hubbard, William 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Hunten, Donald 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Huntsberger, T. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 

Ingersoll, Andrew 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 



Ivanov, A. B. 








MOLA-Based 


Landing 


Site 


Characterization - 


27 






Ivlev, R. 








Planning for a Mars 


in situ sample prepa- 


ration and distribution 


(SPAD) 


system 


- 10 









James, Bonnie 

Aerocapture Technology Project 
Overview - 1 1 

NASA Development of Aerocapture 
Technologies - 14 

Jits, Roman Y. 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

Johnson, Andrew E. 

Optical landmark detection for spacecraft 
navigation - 6 

Johnson, Andrew Edie 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

Johnson, D. L. 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

Johnson, Wyatt R. 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 

- 12 



Johnston, M. D. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Jones, Steven M. 

Mars Sample Return without Landing on 
the Surface - 25 

Josselyn, Scott B. 

Optimization of Low Thrust Trajectories 
With Terminal Aerocapture - 11 

Jurewicz, A. J. G. 

Mars Sample Return without Landing on 
the Surface - 25 

Justus, C. G. 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Atmospheric Models for Aerocapture - 4 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

Engineering-Level Model Atmospheres 
for Titan & Neptune - 13 

Kass, David M, 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Keating, G. M. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Keller, H. U. 

After the Mars Polar Lander: Where to 
Next? - 26 

Keller, Vernon W. 

Atmospheric Models for Aerocapture 
Systems Studies - 1 

Atmospheric Models for Aerocapture - 4 

Atmospheric Models for Aeroentry and 
Aeroassist - 1 

Connecting Atmospheric Science and At- 
mospheric Models for Aerocaptured Mis- 
sions to Titan and the Outer Planets - 5 

Kennedy, Brian M. 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Kerridge, Stuart 

Daily repeat-groundtrack Mars orbits - 7 

Kia, T. 

Micro Navigator - 23 

Kim, S. S. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 

Klarquist, William N. 

Precision Terminal Guidance for a Mars 
Lander - 25 

Knocke, Philip C. 

Mars Exploration Rovers Entry, Descent, 
and Landing Trajectory Analysis - 3 



Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Lam-Trong, Th. 

The Stakes of the Aerocapture for Mis- 
sions to Mars - 25 

Laub, B. 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 

Lee, B. G. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Lee, S. W. 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

Lee, Wayne J. 

Entry descent, and landing scenario for 
the Mars exploration Rover mission - 7 

Leigh, Dennis 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Leshin, L. A. 

After the Mars Polar Lander: Where to 
Next? - 26 

Lewis, Jake 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Liechty, Derek S. 

Control Surface and Afterbody Experi- 
mental Aeroheating for a Proposed Mars 
Smart Lander Aeroshell - 17 

Liever, Peter A. 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Lindstrom, D. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Lingard, Steve J. 

CFD Prediction of the BEAGLE 2 Mars 
Probe Aerodynamic Database - 19 

Lockwood, Mary Kae 

Neptune Aerocapture Systems Analysis 

- 3 

Longuski, James M. 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 

- 12 

Lorenzoni, L. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Lowrie, James W. 

Precision Navigation for a Mars Airplane 

- 27 

Precision Terminal Guidance for a Mars 
Lander - 25 



C-3 



Lyne, J. E. 

Earth Return Aerocapture for 
TransHab/Ellipsled Vehicle - 28 



the 



Lyons, Daniel T. 

Aerobraking at Venus and Mars: A Com- 
parison of the Magellan and Mars Global 
Surveyor Aerobraking Phases - 30 

Pitch control during autonomous aero- 
braking for near-term Mars exploration 

- 12 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Mahaffy, P. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Masciarelli, James P. 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Mase, Robert A. 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

Matthies, Larry H. 

Optical landmark detection for spacecraft 
navigation - 6 

Precise Image-Based Motion Estimation 
for Autonomous Small Body Exploration 

- 30 

May, R. D. 

After the Mars Polar Lander: Where to 
Next? - 26 



McEwen, A. 

The Martian Oasis Detector - 



26 



McNamara, K. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

McRonald, Angus D. 

A Light-Weight Inflatable Hypersonic 
Drag Device for Planetary Entry - 30 

Miller, Kevin L. 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Miller, S. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Millerr, J. H. 

Science and Engineering Potential of an 
Icy Moon Lander - 13 

Moersch, J. E. 

Remote Sensing of Evaporite Minerals in 
Bad water Basin, Death Valley, at Varying 
Spatial Scales and in Different Spectral 
Regions - 22 



NASA Development of Aerocapture 
Technologies - 14 

Morgan, G. 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Morse, Andy 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Muirhead, Brian K. 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

Munk, Michelle M. 

Aeroassist Technology Planning for 
Exploration - 4 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

Munk, Michelle 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Aerocapture Technology Project 
Overview - 11 

NASA Development of Aerocapture 
Technologies - 14 

Muth, W. D. 

Earth Return Aerocapture for the 
TransHab/Ellipsled Vehicle - 28 

Neelon, Joseph 

Daily repeat-groundtrack Mars orbits - 7 



Nilsen, Erik 

Europa Lander 



13 



Noreen, Gary 

Daily repeat-groundtrack Mars orbits - 7 



Oberto, Robert 

Europa Lander 



13 



Moon, Steve 

Aerocapture Technology 
Overview - 1 1 



Project 



OConnell, Tod F. 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Olds, John 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

Olejniczak, Joseph 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Oleson, Steven 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

Olson, Clark F. 

Optical landmark detection for spacecraft 
navigation - 6 



Paige, D. A. 

After the Mars Polar Lander: Where to 
Next? - 26 

Papanastassiou, D. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 

- 10 

Parker, Timothy J. 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Parmar, Devendra S. 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Partridge, Harry 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Patterson, Michael 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

Perot, Etienne 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Pillinger, Colin T. 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Portock, B. M. 

The Strategy for the Second Phase of 
Aerobraking Mars Global Surveyor - 31 

Powell, Richard W. 

Aeroassist Technology Planning for 
Exploration - 4 

Entry trajectory and atmosphere recon- 
struction methodologies for the mars ex- 
ploration rover mission - 7 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

Powell, Richard 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Prabhu, Dinesh 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Prabhu, Ramadas K. 

Experimental Hypersonic Aerodynamic 
Characteristics of the 2001 Mars Sur- 
veyor Precision Lander with Flap - 18 

Praine, Ian 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Queen, Eric M. 

Angle-of-Attack-Modulated Terminal 
Point Control for Neptune Aerocapture 

- 9 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 

- 9 



C-4 



Mars Smart Lander Parachute Simula- 
tion Model - 19 

Queen, Eric 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Raiszadeh, Behzad 

Mars Exploration Rover Terminal De- 
scent Mission Modeling and Simulation 
- 9 

Raiszadeh, Ben 

Mars Smart Lander Parachute Simula- 
tion Model - 19 

Multibody Parachute Flight Simulations 
for Planetary Entry Trajectories Using 
'Equilibrium Points' - 6 



Rice, J. 

The Martian Oasis Detector - 



26 



Rousseau, Stephane 

Aerocapture Guidance Algorithm Com- 
parison Campaign - 18 

Sabahi, Dara 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

Schoenenberger, Mark 

Mars Exploration Rover Six-Degree-Of- 
Freedom Entry Trajectory Analysis - 14 

Schreiber, Jeffrey 

Radioisotope Electric Propulsion for Fast 
Outer Planetary Orbiters - 20 

Schulz, Walkiria 

Study of Orbital Transfers with Aeroas- 
sisted Maneuvers - 21 

Shams, Qamar A. 

Aerothermal Instrumentation Loads To 
Implement Aeroassist Technology in Fu- 
ture Robotic and Human Missions to 
MARS and Other Locations Within the 
Solar System - 20 

Sims, Mark R. 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Skulsky, E. D. 

Mars reconnaissance orbiter design ap- 
proach for high-resolution surface 
imaging - 12 

Smith, John C. 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 

Smith, P. H. 

After the Mars Polar Lander: Where to 
Next? - 26 

The Martian Oasis Detector - 26 

Smythe, William D. 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

Spencer, David A. 

Navigation Strategy for the Mars 2001 
Lander Mission - 29 



Starr, Brett R. 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Stein, Jim 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Stewart, Jenny 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Sture, S. 

Low Velocity Impact Experiments in 
Microgravity - 24 

Sutton, Kenneth 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Takashima, Naruhisa 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Tan-Wang, Grace H. 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

Tartabini, Paul V. 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

The Development and Evaluation of an 
Operational Aerobraking Strategy for the 
Mars 2001 Odyssey Orbiter - 17 

Tolson, R. H. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Approaches to autonomous aerobraking 
at Mars - 15 

Tolson, Robert H. 

Development of a Monte Carlo Mars- 
gram model for 2001 Mars Odyssey 
aerobraking simulations - 15 

Tolson, Robert 

Autonomous Aerobraking at Mars - 16 

tomasko, M. G. 

The Martian Oasis Detector - 26 

Trochman, Bill 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Turner, Andrew E. 

Daily repeat-groundtrack Mars orbits - 7 

Venkatapathy, E. 

Thermal protection system technology 
and facility needs for demanding future 
planetary missions - 8 

Wahl, Beth E. 

Precision Terminal Guidance for a Mars 
Lander - 25 



Walberg, Gerald D. 

Blended control, predictor-corrector guid- 
ance algorithm: An enabling technology 
for Mars aerocapture - 10 

Walberg, Gerald 

An Investigation of Terminal Guidance 
and Control Techniques for a Robotic 
Mars Lander - 28 

Wawrzyniak, Geoffrey G. 

Mars Exploration Rovers Landing Dis- 
persion Analysis - 3 

Way, David 

Uncertainty Optimization Applied to the 
Monte Carlo Analysis of Planetary Entry 
Trajectories - 23 

Weissman, Paul R. 

The Deep Space 4/Champollion Comet 
Rendezvous and Lander Technology 
Demonstration Mission - 28 

Wercinski, Paul 

Aerocapture Technology Development 
Needs for Outer Planet Exploration - 20 

Werner, M. R. 

Application of Accelerometer Data to 
Mars Odyssey Aerobraking and Atmo- 
spheric Modeling - 19 

Westhelle, Carlos H. 

Aerocapture Performance Analysis for a 
Neptune-Triton Exploration Mission - 2 

Wilmoth, Richard G. 

Plume Modeling and Application to Mars 
2001 Odyssey Aerobraking - 16 

Trailing Ballute Aerocapture: Concept 
and Feasibility Assessment - 12 

Wake Closure Characteristics and After- 
body Heating on a Mars Sample Return 
Orbiter- 15 

Wood, G. E. 

Ultra-stable oscillators for planetary entry 
probes - 8 

Wright, Ian P. 

Beagle 2: The Next Exobiology Mission 
to Mars - 21 

Wright, Michael J. 

Preliminary Convective-Radiative Heat- 
ing Environments for a Neptune Aero- 
capture Mission - 1 

Yen, A. S. 

Mars Sample Return without Landing on 
the Surface - 25 

Young, L. A. 

Exploration of Titan Using Vertical Lift 
Aerial Vehicles - 23 

Young, Richard E. 

Summary of the Boulder Entry Probe 
Workshop April 21-22, 2003, Boulder, 
Colorado, USA - 9 

Zimmerman, W. 

Planning for a Mars in situ sample prepa- 
ration and distribution (SPAD) system 
- 10 



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